Attitude And Orbit Control System Research Articles

Magnetometer-Only Kalman Filter Based Algorithms for High Accuracy Spacecraft Attitude Estimation (A Comparative Analysis)

Kalman Filter (KF) based algorithms are the most frequently employed attitude estimation algorithms. Typically, a fully observable system necessitates the use of two distinct sensor types. Therefore, relying on a single sensor, such as a magnetometer, for spacecraft attitude estimation is deemed to be a challenge. The present investigation centers on utilizing magnetometers as the exclusive sensor. Several KF based estimation algorithms have been designed and evaluated to give the designer of spacecraft Attitude and Orbit Control System (AOCS) the choice of a suitable algorithm for his mission based on quantitative measures. These algorithms are capable of effectively addressing nonlinearity in both process and measurement models. The algorithms under examination encompass the Extended Kalman Filter (EKF), Sequential Extended Kalman Filter (SEKF), Pseudo Linear Kalman Filter (PSELIKA), Unscented Kalman Filter (USKF), and Derivative Free Extended Kalman Filter (DFEKF). The comparison of the distinct algorithms hinges on key performance metrics, such as estimation error for each axis, computation time, and convergence rate. The resulting algorithms provide numerous benefits, such as diverse levels of high estimation accuracy (with estimation errors ranging from 0.014o to 0.14o), varying computational demands (execution time ranges from 0.0536s to 0.0584s), and the capability to converge despite large initial attitude estimation errors (which reached 170o). These properties render the algorithms appropriate for utilization by spacecraft designers in all operational modes, supplying high-precision attitude estimations better than (0.5o) despite high magnetometer noise levels, which reached (200 nT).

Visual servoing for coordinated precise attitude control in the TOM small satellite formation

Formations of multiple satellites observe target areas on the Earth’s surface from different perspectives, enabling construction of 3D surface maps by photogrammetric methods. The main objective of TOM (Telematics Earth Observation Mission) is to monitor ash clouds during Volcanic Eruptions with a satellite formation consisting of three CubeSats. TOM is Bavaria’s contribution to the multi-national Telematics International Mission (TIM), composed of 7 satellites by partners from 4 continents. To achieve sufficient overlap of the images for the usage of photogrammetric methods, very precise and coordinated attitude control of the satellites is necessary. This paper presents a vision-based attitude control approach using image data to calculate the control input, also referred to as visual servoing. Thus, the cameras and data processing capacity onboard the CubeSats will be used for both image acquisition and attitude determination, eliminating the need for additional precise attitude sensors like star trackers. To perform visual servoing, highly distinctive features are identified and tracked in the images. The control input for reaction wheels is calculated from the movement of the features between consecutive images. To ensure that the images captured by the different satellites overlap sufficiently for photogrammetric processing, the satellites track the same target area based on feature data exchange directly between the satellites. Simulations of realistic scenarios based on the TOM project show the applicability of the presented attitude control approach. In a first step, a software-based Earth observation simulator generates the acquired images, demonstrating feasibility to compute and track features in satellite images to compute the AOCS (Attitude and Orbit Control System) input. In the second step, hardware-in-the-loop (HiL) simulations allow to realistically characterize the performance of this approach. This HiL environment utilizes two high-precision/high-dynamics turntables as 3-axis attitude simulators, where the satellites with their cameras are mounted, while the observation target is mounted on a mobile robot to simulate the relative orbit motion.

Development of the methodological approaches for the attitude control system of the Earth remote sensing satellite in the conditions of the onboard equipment partial failures

The spacecraft controllability of the angular motion is possible only with operability of the attitude and orbit control system (AOCS) of the spacecraft, sensors, actuators and the spacecraft power system. However, there is a rather significant probability of failure of this equipment during the operation of the spacecraft. This is especially observed after half of the spacecraft's lifetime or because of emergency situations. There is a problem which is connected with providing the maximum performance of the AOCS in case of partial failures of their actuators (reaction wheels (RW), magnetorquer rods (MGTR), etc.).
 Thus, the purpose of this work is the development and synthesis of special algorithms for spacecraft angular motion control in the emergency situations which are connected with RWs partial failures and restrictions of onboard electricity consumption. The approach of synthesis of this control algorithms is based on using mobile control methods which allow to reserve RWs by MGTRs. There are different variants of control loops depending on MGTRs turning on combinations. There were proposed two types of control switching functions: time-periodic and switching by deviation. Also was proposed a methodology of controller synthesis using these switching functions.
 Using this methodology and computer simulation, it was shown the possibility of providing angular nadir orientation and stabilization of the spacecraft with maximum 1−1.5 deg error in case of time-periodic switching functions implementation. Switching by deviation allows to reduce onboard electricity consumption for 25−30 % comparing with using time-periodic switching. However, the accuracy of stabilization significantly lower in case of switching by deviation. Considering these estimates, the corresponding methodological recommendations were formulated for use switching functions depending on emergency

Fault Tolerant Attitude and Orbit Determination System for Small Satellite Platforms

Small satellite platforms are experiencing increasing interest from the space community, because of the reduced cost and the performance available with current technologies. In particular, the hardware composing the attitude and orbit control system (AOCS) has reached a strong maturity level, and the dimensions of the components allow redundant sets of sensors and actuators. Thus, the software shall be capable of managing these redundancies with a fault tolerant structure. This paper presents an attitude and orbit determination system (AODS) architecture, with embedded failure detection and isolation functions, and autonomous redundant component management and reconfiguration for basic failure recovery. The system design and implementation has been sized for small satellite platforms, characterized by limited computing capacities, and reduced autonomy level. The discussion describes the system architecture, with particular emphasis on the failure detection and isolation blocks at the component level. The set of functions managing failure detection at system level is also described in the paper. The proposed system is capable of reconfiguring and autonomously recalibrating after various failures had occurred. Attention is also dedicated to the achieved performance, satisfying stringent requirements for a small satellite platform. In these regards, the simulation results used to verify the performance of the proposed system at the model-in-the-loop (MIL) level are also reported.

Advanced probabilistic [formula omitted]-analysis techniques for AOCS validation

Monte-Carlo simulations play a key role in the current Attitude and Orbit Control Systems (AOCS) Verification and Validation (V&V) process, but it is generally time-consuming and it may fail in detecting worst-case configurations, especially in the presence of rare events. In such a case, μ-analysis offers a nice alternative, although it cannot measure the probability of occurrence of the identified worst-cases, which can invalidate a control system on the basis of unlikely events. Probabilistic μ-analysis was introduced in this context 20 years ago to bridge the gap between the two techniques, but until recently no practical tools were available. This paper summarizes recent advances on this topic with a particular emphasis on practical applications to space systems. More precisely, the proposed technique is applied to evaluate AOCS controllers in the context of a challenging high accuracy satellite pointing control problem. The way the proposed tools can be integrated into the traditional AOCS V&V process and used to tighten the V&V analysis gap is also highlighted.

Numerical study of the switching mechanism of a jet valve using the meshless method

This study numerically investigates fluid dynamics of a jet flow at supersonic speed. The meshless method and the overlapping point cloud method are used to handle the moving boundary problems. The interaction between the jet flow and a moving ball-shaped plug is numerically solved, which has been rarely done in the published literature. The switching mechanism of a novel designed jet valve in an attitude and orbit control system (AOCS) is analyzed. It is found out that applied pressure to the control inlets of the jet valve must be high enough in order to successfully drive the plug to move and subsequently change the force direction acting on the jet valve. Then the switching mechanism of AOCS can be triggered. The initial fluid condition also plays a vital role and it significantly influences the response time of the switch. This study explores the underlying physics of the jet flow on its deflection, wall attachment, and interaction with the ball-shaped plug. It contributes to the optimization design of the jet valve in the AOCS with a fast and efficient response.

Optimal Factorization of the State-Dependent Riccati Equation Technique in a Satellite Attitude and Orbit Control System

The satellite attitude and orbit control system (AOCS) can be designed with success by linear control theory if the satellite has slow angular motions and small attitude maneuver. However, for large and fast maneuvers, the linearized models are not able to represent all the perturbations due to the effects of the nonlinear terms present in the dynamics and in the actuators (e.g., saturation). Therefore, in such cases, it is expected that nonlinear control techniques yield better performance than the linear control techniques. One candidate technique for the design of AOCS control law under a large maneuver is the State-Dependent Riccati Equation (SDRE). SDRE entails factorization (that is, parameterization) of the nonlinear dynamics into the state vector and the product of a matrix-valued function that depends on the state itself. In doing so, SDRE brings the nonlinear system to a (nonunique) linear structure having state-dependent coefficient (SDC) matrices and then it minimizes a nonlinear performance index having a quadratic-like structure. The nonuniqueness of the SDC matrices creates extra degrees of freedom, which can be used to enhance controller performance, however, it poses challenges since not all SDC matrices fulfill the SDRE requirements. Moreover, regarding the satellite's kinematics, there is a plethora of options, e.g., Euler angles, Gibbs vector, modified Rodrigues parameters (MRPs), quaternions, etc. Once again, some kinematics formulation of the AOCS do not fulfill the SDRE requirements. In this paper, we evaluate the factorization options (SDC matrices) for the AOCS exploring the requirements of the SDRE technique. Considering a Brazilian National Institute for Space Research (INPE) typical mission, in which the AOCS must stabilize a satellite in three-axis, the application of the SDRE technique equipped with the optimal SDC matrices can yield gains in the missions. The initial results show that MRPs for kinematics provides an optimal SDC matrix.

Coordinative coupled attitude and orbit control for satellite formation with multiple uncertainties and actuator saturation

This paper studies the coordinative coupled attitude and orbit control system (AOCS) of a satellite formation composed of one microsatellite and two 3U CubeSats. It is described as a highly nonlinear and coupled system subject to actuator saturation and multiple uncertainties including parametric perturbations and external disturbances. To address this issue, the centralized architecture and leader-following method are adopted in the formation firstly. The closed-loop system of each satellite member is divided into orbit and attitude control loops respectively. The coupling relationship between orbit and attitude is introduced as a constraint of the control law. Then a robust controller consisting of a nominal controller, a robust compensator and an advanced proportional–derivative (APD) controller is proposed to achieve high-accuracy formation control. The robust compensator is constructed from an auxiliary system and a finite-time observer to compensate the influence of an equivalent lumped disturbance. The APD controller is developed to achieve desired performances under actuator saturation. Stability analysis indicates that both relative position errors and attitude tracking errors can converge asymptotically. Finally, numerical simulations are carried out to verify the effectiveness of proposed control law in both relatively-stationary and fly-around scenarios.

Study on Model-Based System Engineering in Satellite Layout

For satellite layout, it is necessary to collate manually in order to obtain accurate product technical requirements information. It is difficult to track the technical status information. It is proposed that the application of model-based system engineering (MBSE) in the cooperative development of multi-disciplinary and multisectoral satellite layout was an effective way to solve these problems. Based on the four typical processes of layout design, this paper studied the requirements to occupancy design, electromechanical joint design, integration of Attitude and Orbit Control System (AOCS) equipment layout, definition and display of equipment expansion, and give the verification design of star sensor layout. The results show that this method is beneficial to the collaboration of all departments, and improve the design ability.

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.

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.

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.

Space Debris Removal using a Tether: A Model

Focused on the removal of space debris, this paper studies the modeling of a target satellite connected to a chaser satellite by a tether. All dynamic couplings between the flexible and rigid modes of the satellites are accounted for. The tether attached to both satellites is modeled as a massless spring when stretched and non-existent when compressed. The objective of this work is to model the behaviour of the three bodies: chaser, tether, and target in the orbital reference frame for arbitrary initial conditions of the target satellite. Simulations of the whole system including the chaser Attitude and Orbit Control System (AOCS) are performed to evaluate its ability to damp the debris tumbling motion and to tow the debris.

Editorial: International ESA Conference on Guidance, Navigation & Control Systems 2014

miniaturisation efforts as well as techniques for advanced estimation, control and verification. Two very high accuracy missions, namely Euclid and GOES-R, are presented in the first two papers. They provide a good illustration of the overall engineering process, which includes mission objectives flow down to Attitude and Orbit Control System (AOCS) requirements, sensors and actuators selection, control design analysis and performance verification. Euclid is a medium-class mission of the ESA Cosmic Vision 2015–2025 plan. Operating around the L2 Lagrangian point of the Sun–Earth system, it will seek to understand the origin of the universe’s accelerating expansion. GOES-R will be the first of the next-generation geostationary weather satellites, managed by NOAA and NASA. Both missions have unprecedented pointing performance and image quality requirements, which require innovative pointing architecture and implementation. Planetary exploration missions are also highly inspirational and demanding in terms of GNC-enabling techniques. Mars exploration is addressed in three papers: firstly, with the Exomars 2016 mission and the Orbiter GNC implementing aerobraking strategies; secondly, with the in-flight experience gained from the Entry, Descent and Landing of the Mars Science Laboratory; and lastly with vision-based autonomous Rendezvous and Capture technologies under development for future Mars sample and return missions. Lunar exploration is then covered by a paper on autonomous landing with hazard detection and avoidance techniques. New challenges related to space debris mitigation are presented in two papers on active debris removal. The first is a general survey of the GNC-related needs, followed by a student paper on the combined control of a chaser and its robotic arm. This special issue of the CEAS Space Journal has been prepared from a selection of the most interesting presentations delivered throughout the different sessions of the 9th International ESA Conference on Guidance, Navigation & Control Systems. Every three years, this conference offers a unique window into the most recent experiences and innovations of the international GNC community, as well as a detailed insight into the demanding requirements of future missions. The latest conference, held in June 2014 in Porto, was attended by more than 250 GNC engineers, managers and decision makers from around the globe. A broad range of GNC-related topics was covered by presentations of the highest calibre. Thanks to the authors who accepted to submit their papers to the CEAS Space Journal, and thanks to the reviewers, we are happy to provide a representative and up-to-date survey covering challenging missions and associated GNC developments, new sensors including

A Fault-Tolerant and Anti-Disturbance Control Based on Sliding Mode Iterative Learning Law for Attitude of Satellite

To cope with the problem of the degradation of actuation effectiveness caused by actuator deflection or fault in the attitude and orbit control system (AOCS) of spacecraft on-orbit, an attitude fault-tolerant and anti-disturbance control scheme is proposed based on a sliding mode iterative learning law, in which the pseudo control input is applied to design the sliding mode controller to ensure the AOCS tracks a reference trajectory precisely after some fault occurred; By analyzing the Lyapunov stability, a novel adaptive iterative learning law is developed, which in term of the tracking error, determine some parameters in controller on-line to address the actuator failure and external disturbance. Numerical simulation experiments show that the fault-tolerant and anti-disturbance controller can ameliorate actuation malfunction and compensate the influence of external disturbance effectively.

A symbolic model checking approach to verifying satellite onboard software

This paper discusses the use of symbolic model checking technology to verify the design of an embedded satellite software control system called the attitude and orbit control system (AOCS). This system is mission critical because it is responsible for maintaining the attitude of the satellite and for performing fault detection, isolation, and recovery decisions. An executable AOCS implementation by Space Systems Finland has been provided in Ada source code form, and we use the input language of the symbolic model checker NuSMV 2 to model the implementation at a detailed level. We describe the modeling techniques and abstractions used to alleviate the state space explosion due to the handling of timers and the large number of system components controlled by the AOCS. The required behavior has been specified as extended state machine diagrams and translated to temporal logic properties. Besides well-known LTL and CTL model checking algorithms, we adapt a previously unexplored form of the liveness-to-safety approach to the problem. The latter new technique turns out to successfully prove all desired properties of the system, outperforming both the LTL and CTL implementations of NuSMV 2.

Flat Stellar Dynamic Filter For Gyroless Attitude Estimation

In this paper, we present an innovative attitude estimation filter based on differential flatness of the satellite dynamics with Rodriguez Modified Parameters for star tracker only Attitude and Orbit Control System (AOCS) architectures. The main characteristic of this estimator is to be able to maintain the satellite Absolute Pointing Error (APE) within safe limits, avoiding safe hold mode recovery in case of long duration loss of measurements (attitude control with no measurements for 1000s). This filter has been selected for in flight testing on PARASOL CNES microsatellite, and will be experimented by May 2013.

Developing mode-rich satellite software by refinement in Event-B

One of the guarantees that the designers of on-board satellite systems need to provide, so as to ensure their dependability, is that the mode transition scheme is implemented correctly, i.e. that the states of system components are consistent with the global system mode. There is still, however, a lack of scalable approaches to developing and verifying systems with complex mode transitions. This paper presents an approach to the formal development of mode-rich systems by refinement in Event-B. We formalise the concepts of modes and mode transitions as well as deriving specification and refinement patterns which support correct-by-construction system development. The proposed approach is validated by a formal development of the Attitude and Orbit Control System (AOCS) undertaken within the ICT DEPLOY project. The experience gained in the course of developing such a complex industrial system as AOCS, shows that Event-B refinement provides the engineers with a scalable formal technique. Moreover, the case study has demonstrated that Event-B can facilitate formal development of mode-rich systems and, in particular, proof-based verification of their mode consistency.

Characteristics and accuracies of the GRACE inter-satellite pointing

For almost 10years, the Gravity Recovery and Climate Experiment (GRACE) has provided information about the Earth gravity field with unprecedented accuracy. Efforts are ongoing to approach the GRACE baseline accuracy as there still remains an order of magnitude between the present error level of the gravity field solutions and the GRACE baseline. At the current level of accuracy, thorough investigation of sensor related effects is necessary as they are one of the potential contributors to the error budget. In the science mode operations, the twin satellites are kept precisely pointed with their KBR antennas towards each other. It is the task of the onboard attitude and orbit control system (AOCS) to keep the satellites in the required formation. We analyzed long time series of the inter-satellite pointing variations as they reflect the AOCS performance and characteristics. We present significant systematic effects in the inter-satellite pointing and discuss their possible sources. Prominent features are especially related to the magnetic torquer characteristics, star cameras’ performance and KBR antenna calibration parameters. The relation between the magnetic torquer attitude control and the Earth magnetic field, impact of the different performance of the two star camera heads on the attitude control and the features due to uncertainties in the calibration parameters relating the star camera frame to K-frame are discussed in detail. Proper understanding of these effects will help to reduce their impact on the science data and subsequently increase the accuracy of the gravity field solutions. Moreover, understanding the complexity of the onboard system is essential not only for increasing the accuracy of the GRACE data but also for the development of the future gravity field satellite missions.

FDI(R) for satellites: How to deal with high availability and robustness in the space domain?

FDI(R) for satellites: How to deal with high availability and robustness in the space domain?The European leader for satellite systems and at the forefront of orbital infrastructures, Thales Alenia Space, is a joint venture between Thales (67%) and Finmeccanica (33%) and forms with Telespazio a Space Alliance. Thales Alenia Space is a worldwide reference in telecoms, radar and optical Earth observation, defence and security, navigation and science. It has 11 industrial sites in 4 European countries (France, Italy, Spain and Belgium) with over 7200 employees worldwide. Satellite evolution and the wish to design more autonomous missions imply the enhancement of the satellite architecture and special attention paid to fault management (i.e., Fault Detection, Isolation and Recovery, or FDIR, in space). Nevertheless, the constraints on FDIR techniques and strategies remain the same as for standard missions: robustness, reactive detection, quick isolation/identification and validation. This paper gives an introduction to Fault Tolerance (FT) in the space domain and some principles for the coming FT architectures. The current context of FDIR is presented by describing the approach implemented on telecommunication satellites and, more precisely, on one of the most FDIR sensible subsystems: the AOCS (Attitude and Orbit Control System). Following the current state of FDIR in the space domain, some perspectives are given such as a centralized distributed FDIR strategy for the next generation of autonomous satellites as well as some research tracks and hybrid diagnosis.