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Simulations are made to compare the new filter with the traditional EKF. The results indicate that under same conditions, compared with EKF, UKF has faster convergence speed, higher filtering accuracy and more stable estimation performance. The satellite has been designed and built, by students Full Text Available Attitude control of artificial satellites is dependent on information provided by its attitude determination process.

This paper presents the implementation and tests of a fully self-contained algorithm for the attitude determination using magnetometers and accelerometers, for application on a satellite simulator based on frictionless air bearing tables.

However, it is known that magnetometers and accelerometers need to be calibrated so as to allow that measurements are used to their ultimate accuracy. A calibration method is implemented which proves to be essential for improving attitude determination accuracy. For the stepwise real-time attitude determination , it was used the well-known QUEST algorithm which yields quick response with reduced computer resources.

The algorithms are tested and qualified with actual data collected on the streets under controlled situations. For such street runaways, the experiment employs a solid-state magnetoresistive magnetometer and an IMU navigation block consisting of triads of accelerometers and gyros, with MEMS technology. A GPS receiver is used to record positional information. The collected measurements are processed through the developed algorithms, and comparisons are made for attitude determination using calibrated and noncalibrated data.

The results show that the attitude accuracy reaches the requirements for real-time operation for satellite simulator platforms. A semi-physical simulation platform of attitude determination and control system for satellite. Full Text Available A semi-physical simulation platform for attitude determination and control system is proposed to verify the attitude estimator and controller on ground.

A simulation target, a host PC, many attitude sensors, and actuators compose the simulation platform. A three-axes gyroscope, a three-axes magnetometer, a sun sensor, a star tracer, three flywheels, and a Global Positioning System receiver are connected to the simulation target, which formulates the attitude control cycle of a satellite. The simulation models of the attitude determination and control system are described in detail.

Finally, the semi-physical simulation platform is used to demonstrate the availability and rationality of the control scheme of a micro- satellite. Comparing the results between the numerical simulation in Simulink and the semi-physical simulation, the semi-physical simulation platform is available and the control scheme successfully achieves three-axes stabilization. They have improved in the past decade, from relatively low resolution at about 5m to Until August of , there was only one ADS available on the market.

It was the Pumpkin. Satellite single-axis attitude determination based on Automatic Dependent Surveillance - Broadcast signals. The satellite equipped with spaceborne ADS-B system receives the broadcast signals from aircraft and transfers the message to ground stations, so as to extend the coverage area of terrestrial-based ADS-B.

In this work, a novel satellite single-axis attitude determination solution based on the ADS-B receiving system is proposed. This solution utilizes the signal-to-noise ratio SNR measurement of the broadcast signals from aircraft to determine the boresight orientation of the ADS-B receiving antenna fixed on the satellite.

The basic principle of this solution is described. The feasibility study of this new attitude determination solution is implemented, including the link budget and the access analysis. On this basis, the nonlinear least squares estimation based on the Levenberg-Marquardt method is applied to estimate the single-axis orientation. A full digital simulation has been carried out to verify the effectiveness and performance of this solution.

Finally, the corresponding results are processed and presented minutely. An RAE-I satellite description is given, taking into account a dynamics experiment and the attitude sensing system. A computer program for analyzing flexible spacecraft attitude motions is considered, giving attention to the geometry of rod deformation.

The characteristics of observed attitude data are discussed along with an analysis of the main boom root angle, the bending rigidity, and the damper plane angle. Error analysis of satellite attitude determination using a vision-based approach. Improvements in communication and processing technologies have opened the doors to exploit on-board cameras to compute objects' spatial attitude using only the visual information from sequences of remote sensed images.

The strategies and the algorithmic approach used to extract such information affect the estimation accuracy of the three-axis orientation of the object. This work presents a method for analyzing the most relevant error sources, including numerical ones, possible drift effects and their influence on the overall accuracy, referring to vision-based approaches. The method in particular focuses on the analysis of the image registration algorithm, carried out through on-purpose simulations.

The overall accuracy has been assessed on a challenging case study, for which accuracy represents the fundamental requirement. In particular, attitude determination has been analyzed for small satellites , by comparing theoretical findings to metric results from simulations on realistic ground-truth data.

Significant laboratory experiments, using a numerical control unit, have further confirmed the outcome. We believe that our analysis approach, as well as our findings in terms of error characterization, can be useful at proof-of-concept design and planning levels, since they emphasize the main sources of error for visual based approaches employed for satellite attitude estimation. Nevertheless, the approach we present is also of general interest for all the affine applicative domains which require an accurate estimation of three-dimensional orientation parameters i.

Using the global positioning satellite system to determine attitude rates using doppler effects. In the absence of a gyroscope, the attitude and attitude rate of a receiver can be determined using signals received by antennae on the receiver. Based on the signals received by the antennae, the Doppler difference between the signals is calculated. The Doppler difference may then be used to determine the attitude rate.

With signals received from two signal sources by three antennae pairs, the three-dimensional attitude rate is determined. On-orbit real-time magnetometer bias determination for micro- satellites without attitude information. Full Text Available Due to the disadvantages such as complex calculation, low accuracy of estimation, and being non real time in present methods, a new real-time algorithm is developed for on-orbit magnetometer bias determination of micro- satellites without attitude knowledge in this paper.

This method uses the differential value approach. It avoids the impact of quartic nature and uses the iterative method to satisfy real-time applications. Simulation results indicate that the new real-time algorithm is more accurate compared with other methods, which are also tested by an experiment system using real noise data. With the new real-time algorithm, a magnetometer calibration can be taken on-orbit and will reduce the demand for computing power effectively.

Attitude determination for small satellites using GPS signal-to-noise ratio. The design serves as an evaluation testbed for conducting ground based experiments using various computational methods and antenna types to determine the optimum AD accuracy. Raw GPS data is also stored to non-volatile memory for downloading and post analysis. Two low-power microcontrollers are used for processing and to display information on a graphic screen for real-time performance evaluations.

A new parallel inter-processor communication protocol was developed that is faster and uses less power than existing standard protocols. A shorted annular patch SAP antenna was fabricated for the initial ground-based AD experiments with the testbed. Evaluation of geomagnetic field models using magnetometer measurements for satellite attitude determination system at low earth orbits: Case studies. In this study, different geomagnetic field models are compared in order to study the errors resulting from the representation of magnetic fields that affect the satellite attitude system.

The comparisons were made during geomagnetically active and quiet days to see the effects of the geomagnetic storms and sub-storms on the predicted and observed magnetic fields and angles. The angles, in turn, are used to estimate the spacecraft attitude and hence, the differences between model and observations as well as between two models become important to determine and reduce the errors associated with the models under different space environment conditions.

We show that the models differ from the observations even during the geomagnetically quiet times but the associated errors during the geomagnetically active times increase. We find that the T89 model gives closer predictions to the observations, especially during active times and the errors are smaller compared to the IGRF model. For the first time, the geomagnetic models were used to address the effects of the near Earth space environment on the satellite attitude.

Spacecraft Attitude Determination. This thesis describes the development of an attitude determination system for spacecraft based only on magnetic field measurements. The need for such system is motivated by the increased demands for inexpensive, lightweight solutions for small spacecraft. These spacecraft demands full attitude Meeting these objectives with a single vector magnetometer is difficult and requires temporal fusion of data in order to avoid local observability problems.

In order to guaranteed globally nonsingular solutions, quaternions are generally the preferred attitude The quantitative effects of errors in the process and noise statistics are discussed in detail. Visual attitude propagation for small satellites. As electronics become smaller and more capable, it has become possible to conduct meaningful and sophisticated satellite missions in a small form factor.

However, the capability of small satellites and the range of possible applications are limited by the capabilities of several technologies, including attitude determination and control systems. This dissertation evaluates the use of image-based visual attitude propagation as a compliment or alternative to other attitude determination technologies that are suitable for miniature satellites. The concept lies in using miniature cameras to track image features across frames and extracting the underlying rotation.

The problem of visual attitude propagation as a small satellite attitude determination system is addressed from several aspects: related work, algorithm design, hardware and performance evaluation, possible applications, and on-orbit experimentation. These areas of consideration reflect the organization of this dissertation. A "stellar gyroscope" is developed, which is a visual star-based attitude propagator that uses relative motion of stars in an imager's field of view to infer the attitude changes.

The device generates spacecraft relative attitude estimates in three degrees of freedom. Algorithms to perform the star detection, correspondence, and attitude propagation are presented. The Random Sample Consensus RANSAC approach is applied to the correspondence problem to successfully pair stars across frames while mitigating falsepositive and false-negative star detections.

This approach provides tolerance to the noise levels expected in using miniature optics and no baffling, and the noise caused by radiation dose on orbit. The hardware design and algorithms are validated using test images of the night sky. The application of the stellar gyroscope as part of a CubeSat attitude determination and control system is described. The stellar gyroscope is used to augment a MEMS gyroscope attitude propagation. Innovative power management, attitude determination and control tile for CubeSat standard Nano Satellites.

Electric power supply EPS and attitude determination and control subsystem ADCS are the most essential elements of any aerospace mission. So keeping in mind their importance, they have been integrated and developed on a single tile called CubePMT module. Modular power management tiles PMTs are already available in the market but they are less efficient, heavier in weight, consume more power and contain less number of subsystems.

CubePMT is developed on the design approach of AraMiS architecture: a project developed at Politecnico di Torino that provides low cost and higher performance space missions with dimensions larger than CubeSats. The feature of AraMiS design approach is its modularity. These modules can be reused for multiple missions which helps in significant reduction of the overall budget, development and testing time.

One has just to reassemble the required subsystems to achieve the targeted specific mission. Incorporation of star measurements for the determination of orbit and attitude parameters of a geosynchronous satellite : An iterative application of linear regression.

The image coordinates of the star locations are measured and stored. Subsequently, the information is used to determine the attitude , the misalignment angles between the spin axis and the principal axis of the satellite , and the precession rate and direction. This is done for both the 'East' and 'West' operational geosynchronous satellites. This orientation information is then combined with image measurements of earth based landmarks to determine the orbit of each satellite.

The method for determining the orbit is simple. For each landmark measurement one determines a nominal position vector for the satellite by extending a ray from the landmark's position towards the satellite and intersecting the ray with a sphere with center coinciding with the Earth's center and with radius equal to the nominal height for a geosynchronous satellite.

The apparent motion of the satellite around the Earth's center is then approximated with a Keplerian model. In turn the variations of the satellite 's height, as a function of time found by using this model, are used to redetermine the successive satellite positions by again using the Earth based landmark measurements and intersecting rays from these landmarks with the newly determined spheres.

This process is performed iteratively until convergence is achieved. Only three iterations are required. Satellite Attitude Control System Simulator. Full Text Available Future space missions will involve satellites with great autonomy and stringent pointing precision, requiring of the Attitude Control Systems ACS with better performance than before, which is function of the control algorithms implemented on board computers.

The difficulties for developing experimental ACS test is to obtain zero gravity and torque free conditions similar to the SCA operate in space. However, prototypes for control algorithms experimental verification are fundamental for space mission success. This paper presents the parameters estimation such as inertia matrix and position of mass centre of a Satellite Attitude Control System Simulator SACSS, using algorithms based on least square regression and least square recursive methods.

Simulations have shown that both methods have estimated the system parameters with small error. The SACSS platform model will be used to do experimental verification of fundamental aspects of the satellite attitude dynamics and design of different attitude control algorithm. As three axis stabilization is high level technology requiring the proper functioning of various sensors, actuators and control software, many early satellites failed in their initial operation phase because of shortage of solar power generation or inability to realize the initial step of missions because of unexpected attitude control system performance.

These results come from failure to design the satellite attitude determination and control system ADCS appropriately and not considering " satellite survivability. This paper discusses how to realize ADCS while taking satellite survivability into account, based on our experiences of design and in-orbit operations of Hodoyoshi-3 and 4 satellites launched in , which suffered from various component anomalies but could complete their missions.

Passive Magnetic Attitude Control PMAC is capable of aligning a satellite within 5 degrees of the local magnetic field at low resource cost, making it ideal for a small satellite. However, simulation attempts to date have not been able to predict the attitude dynamics at a level sufficient for mission design. Also, some satellites have suffered from degraded performance due to an incomplete understanding of PMAC system design. This dissertation alleviates these issues by discussing the design, inputs, and validation of PMAC systems for small satellites.

After on-orbit calibration of the off-the-shelf magnetometer and photodiodes and an on-orbit fit to the satellite magnetic moment, the MEKF regularly achieves a three sigma attitude uncertainty of 4 degrees or less. CSSWE is found to settle to the magnetic field in seven days, verifying its attitude design requirement. A Helmholtz cage is constructed and used to characterize the CSSWE bar magnet and hysteresis rods both individually and in the flight configuration.

Fitted parameters which govern the magnetic material behavior are used as input to a PMAC dynamics simulation. All components of this simulation are described and defined. Simulation-based dynamics analysis shows that certain initial conditions result in abnormally decreased settling times; these cases may be identified by their dynamic response. The simulation output is compared to the MEKF output; the true dynamics are well modeled and the predicted settling time is found to possess a 20 percent error, a significant improvement over prior simulation.

Satellite recovery - Attitude dynamics of the targets. The problems of categorizing and modeling the attitude dynamics of uncontrolled artificial earth satellites which may be targets in recovery attempts are addressed. Methods of classification presented are based on satellite rotational kinetic energy, rotational angular momentum and orbit and on the type of control present prior to the benign failure of the control system.

The use of approximate analytical solutions and 'exact' numerical solutions to the equations governing satellite attitude motions to predict uncontrolled attitude motion is considered. Analytical and numerical results are presented for the evolution of satellite attitude motions after active control termination. While there are a multitude of ways to determine a satellite 's orientation, very little research has been done on determining if the attitude of a satellite can be determined directly from telemetry The three ACTS control axes are defined, including the means for sensing attitude and determining the pointing errors.

The desired pointing requirements for various modes of control as well as the disturbance torques that oppose the control are identified. Finally, the hardware actuators and control loops utilized to reduce the attitude error are described. A new star tracker concept for satellite attitude determination based on a multi-purpose panoramic camera. This paper presents an innovative algorithm developed for attitude determination of a space platform.

The algorithm exploits images taken from a multi-purpose panoramic camera equipped with hyper-hemispheric lens and used as star tracker. The sensor architecture is also original since state-of-the-art star trackers accurately image as many stars as possible within a narrow- or medium-size field-of-view, while the considered sensor observes an extremely large portion of the celestial sphere but its observation capabilities are limited by the features of the optical system.

The proposed original approach combines algorithmic concepts, like template matching and point cloud registration, inherited from the computer vision and robotic research fields, to carry out star identification. The final aim is to provide a robust and reliable initial attitude solution lost-in-space mode , with a satisfactory accuracy level in view of the multi-purpose functionality of the sensor and considering its limitations in terms of resolution and sensitivity.

Performance evaluation is carried out within a simulation environment in which the panoramic camera operation is realistically reproduced, including perturbations in the imaged star pattern. Confined computer capacity and a limit on electrical power supply were separate obstacles. They demanded computational simplicity and power optimality from the attitude control system. The design of quasi optimal controllers for a real-time implementation It was explained how the equilibria depended on the ratio of the satellite 's moments of inertia.

It was further investigated how to control the attitude , such that the satellite was globally asymptotically stable in the desired Satellite Photometric Error Determination. Payne, Philip J. Castro, Stephen A. The Johnson photometric system is a set of filters in the optical. Attitude stability analyses for small artificial satellites. The objective of this paper is to analyze the stability of the rotational motion of a symmetrical spacecraft, in a circular orbit.

The equilibrium points and regions of stability are established when components of the gravity gradient torque acting on the spacecraft are included in the equations of rotational motion, which are described by the Andoyer's variables. The nonlinear stability of the equilibrium points of the rotational motion is analysed here by the Kovalev-Savchenko theorem.

With the application of the Kovalev-Savchenko theorem, it is possible to verify if they remain stable under the influence of the terms of higher order of the normal Hamiltonian. In this paper, numerical simulations are made for a small hypothetical artificial satellite. Several stable equilibrium points were determined and regions around these points have been established by variations in the orbital inclination and in the spacecraft principal moment of inertia.

The present analysis can directly contribute in the maintenance of the spacecraft's attitude. Statistical Attitude Determination. All spacecraft require attitude determination at some level of accuracy. This can be a very coarse requirement of tens of degrees, in order to point solar arrays at the sun, or a very fine requirement in the milliarcsecond range, as required by Hubble Space Telescope.

A toolbox of attitude determination methods, applicable across this wide range, has been developed over the years. There have been many advances in the thirty years since the publication of Reference, but the fundamentals remain the same. One significant change is that onboard attitude determination has largely superseded ground-based attitude determination , due to the greatly increased power of onboard computers.

The availability of relatively inexpensive radiation-hardened microprocessors has led to the development of "smart" sensors, with autonomous star trackers being the first spacecraft application. Another new development is attitude determination using interferometry of radio signals from the Global Positioning System GPS constellation.

This article reviews both the classic material and these newer developments at approximately the level of, with emphasis on. We discuss both "single frame" methods that are based on measurements taken at a single point in time, and sequential methods that use information about spacecraft dynamics to combine the information from a time series of measurements.

Introducing the state variable function of GPS attitude determination algorithm in SAR satellite by means of kinematic vector and describing the observation function by the GPS wide-band carrier phase, the paper uses the Kalman filter algorithm to obtian the attitude variables of SAR satellite. Compared the simulation results of Kalman filter algorithm with the least square algorithm and explicit solution, it is indicated that the Kalman filter algorithm is the best.

Full Text Available This article has discussed the development of a three-axis attitude digital controller for an artificial satellite using a digital signal processor. The main motivation of this study is the attitude control system of the satellite Multi-Mission Platform, developed by the Brazilian National Institute for Space Research for application in different sort of missions.

The controller design was based on the theory of the Linear Quadratic Gaussian Regulator, synthesized from the linearized model of the motion of the satellite , i. In the first stage of the project development, a system controller for continuous time was studied with the aim of testing the adequacy of the adopted control.

Star trackers for attitude determination. One problem comes to all spacecrafts using vector information. That is the problem of determining the attitude. This paper describes how the area of attitude determination instruments has evolved from simple pointing devices into the latest technology, which determines the attitude by utilizing The instruments are called star trackers and they are capable of determining the attitude with an accuracy better than 1 arcsecond.

The concept of the star tracker is explained. The obtainable accuracy is calculated, the numbers of stars to be included Finally the commercial market for star trackers is discussed Touchless attitude correction for satellite with constant magnetic moment. Rescue of satellite with attitude fault is of great value. Satellite with improper injection attitude may lose contact with ground as the antenna points to the wrong direction, or encounter energy problems as solar arrays are not facing the sun.

Improper uploaded command may set the attitude out of control, exemplified by Japanese Hitomi spacecraft. In engineering practice, traditional physical contact approaches have been applied, yet with a potential risk of collision and a lack of versatility since the mechanical systems are mission-specific. This paper puts forward a touchless attitude correction approach, in which three satellites are considered, one having constant dipole and two having magnetic coils to control attitude of the first.

Particular correction configurations are designed and analyzed to maintain the target's orbit during the attitude correction process. A reference coordinate system is introduced to simplify the control process and avoid the singular value problem of Euler angles. Based on the spherical triangle basic relations, the accurate varying geomagnetic field is considered in the attitude dynamic mode.

Sliding mode control method is utilized to design the correction law. Finally, numerical simulation is conducted to verify the theoretical derivation. It can be safely concluded that the no-contact attitude correction approach for the satellite with uniaxial constant magnetic moment is feasible and potentially applicable to on-orbit operations. Integrated orbit and attitude hardware-in-the-loop simulations for autonomous satellite formation flying. Development and experiment of an integrated orbit and attitude hardware-in-the-loop HIL simulator for autonomous satellite formation flying are presented.

The integrated simulator system consists of an orbit HIL simulator for orbit determination and control, and an attitude HIL simulator for attitude determination and control. The integrated simulator involves four processes orbit determination , orbit control, attitude determination , and attitude control , which interact with each other in the same way as actual flight processes do. Orbit determination is conducted by a relative navigation algorithm using double-difference GPS measurements based on the extended Kalman filter EKF.

Orbit control is performed by a state-dependent Riccati equation SDRE technique that is utilized as a nonlinear controller for the formation control problem. Attitude is determined from an attitude heading reference system AHRS sensor, and a proportional-derivative PD feedback controller is used to control the attitude HIL simulator using three momentum wheel assemblies. Integrated orbit and attitude simulations are performed for a formation reconfiguration scenario. By performing the four processes adequately, the desired formation reconfiguration from a baseline of m was achieved with meter-level position error and millimeter-level relative position navigation.

This HIL simulation demonstrates the performance of the integrated HIL simulator and the feasibility of the applied algorithms in a real-time environment. Furthermore, the integrated HIL simulator system developed in the current study can be used as a ground-based testing environment to reproduce possible actual satellite formation operations. Sensor fault detection and recovery in satellite attitude control.

This paper proposes an integrated sensor fault detection and recovery for the satellite attitude control system. By introducing a nonlinear observer, the healthy sensor measurements are provided. Considering attitude dynamics and kinematic, a novel observer is developed to detect the fault in angular rate as well as attitude sensors individually or simultaneously.

There is no limit on type and configuration of attitude sensors. By designing a state feedback based control signal and Lyapunov stability criterion, the uniformly ultimately boundedness of tracking errors in the presence of sensor faults is guaranteed. Finally, simulation results are presented to illustrate the performance of the integrated scheme.

This has been a serious obstacle for using magnetorquer based control for three-axis attitude control. This paper deals with three-axis stabilization of a low earth orbit satellite. The problem of controlling A three dimensional sliding manifold is proposed, and it is shown that the satellite motion on the sliding manifold is asymptotically stable Low-power attitude determination for magnetometry planetary missions.

This work covers the subject of orientation or attitude in space and on the surface of a planet. Different attitude sensor technologies have been investigated with emphasis on very low power consumption and mass. In addition robust methods for attitude determination have been covered again A true low-power attitude sensor using the Anisotropic Magneto Resistor effect have been designed to late prototype state.

Two prototypes of the AMR magnetometer have been built. One of the prototypes has an analog output and the second Different attitude representations such as orthogonal matrices, Euler angles and quaternions are presented. Also methods for attitude determination of a sensor platform with more than one vector Performance comparison of attitude determination , attitude estimation, and nonlinear observers algorithms.

This paper presents a brief synthesis and useful performance analysis of different attitude filtering algorithms attitude determination algorithms, attitude estimation algorithms, and nonlinear observers applied to Low Earth Orbit Satellite in terms of accuracy, convergence time, amount of memory, and computation time.

This latter is calculated in two ways, using a personal computer and also using On-board computer OBC that is being used in many SSTL Earth observation missions. The use of this comparative study could be an aided design tool to the designer to choose from an attitude determination or attitude estimation or attitude observer algorithms. The simulation results clearly indicate that the nonlinear Observer is the more logical choice.

Hughey, Raymond H. Cost, size, weight and power requirements, on the other hand, impose selecting relative simple sensors and actuators which leads to an attitude control requirement of less than 1 degree. This precision is obtained by a combination of magnetorquers and momentum wheels Attitude Determination and Control Systems. In the year , Galveston, Texas, was a bustling community of approximately 40, people.

The former capital of the Republic of Texas remained a trade center for the state and was one of the largest cotton ports in the United States. On September 8 of that year, however, a powerful hurricane struck Galveston island, tearing the Weather Bureau wind gauge away as the winds exceeded mph and bringing a storm surge that flooded the entire city.

The worst natural disaster in United States history even today the hurricane caused the deaths of between and people. Critical in the events that led to such a terrible loss of life was the lack of precise knowledge of the strength of the storm before it hit.

In , Hurricane Ike, the third costliest hurricane ever to hit the United States coast, traveled through the Gulf of Mexico. Ike was gigantic, and the devastation in its path included the Turk and Caicos Islands, Haiti, and huge swaths of the coast of the Gulf of Mexico. Once again, Galveston, now a city of nearly 60,, took the direct hit as Ike came ashore. Almost people in the Caribbean and the United States lost their lives; a tragedy to be sure, but far less deadly than the storm. This time, people were prepared, having received excellent warning from the GOES satellite network.

The Geostationary Operational Environmental Satellites have been a continuous monitor of the world's weather since , and they have since been joined by other Earth-observing satellites. This weather surveillance to which so many now owe their lives is possible in part because of the ability to point accurately and steadily at the Earth below. The importance of accurately pointing spacecraft to our daily lives is pervasive, yet somehow escapes the notice of most people. But the example of the lives saved from Hurricane Ike as compared to the storm is something no one should ignore.

In this section, we will summarize the processes and technologies used in. Development of a hardware-in-loop attitude control simulator for a CubeSat satellite. Attitude control is an important part in satellite on-orbit operation. It greatly affects the performance of satellites. Testing of an attitude determination and control subsystem ADCS is very challenging since it might require attitude dynamics and space environment in the orbit. The simulator consists of a numerical simulation part, a hardware part, and a HIL interface hardware unit.

The hardware part is the real ADCS board of the satellite. Then, based on this information, the HIL interface hardware generates I2C signals mimicking the signals of the on-board rate-gyros and magnetometers and consequently outputs the signals to the ADCS board. The ADCS board reads the rate-gyro and magnetometer signals, calculates control signals, and drives the attitude actuators which are three magnetic torquers MTQs.

The responses of the MTQs sensed by a separated magnetometer are feedback to the numerical simulation part completing the HIL simulation loop. Experimental studies are conducted to demonstrate the feasibility and effectiveness of the simulator. Full Text Available Attitude data, which is the important data strongly correlated with the geometric accuracy of optical remote sensing satellite images, are generally obtained using a real-time Extended Kalman Filter EKF with star-tracker and gyro data for current high-resolution satellites , such as Orb-view, IKONOS, Quickbird,Pleiades, and ZY Moreover, this method makes full use of the collected data in the fixed-interval and computational resources on the ground, and it determines optimal attitude results by forward-backward filtering and weighted smoothing with the raw star-tracker and gyro data collected for a fixed period.

In this study, the principle and implementation of the proposed method are described. The post-processed attitude was compared with the on-board attitude , and the absolute accuracy was evaluated by the two methods. One method compares the positioning accuracy of the object space coordinates with the post-processed and on-board attitude data without using ground control points GCPs. The other method compares the tie-point residuals of the image coordinates after a free net adjustment.

In addition, the internal and external parameters of the camera were accurately calibrated before use for an objective evaluation of the attitude accuracy. The experimental results reveal that the accuracy of the post-processed attitude is superior to the accuracy of the on-board processed attitude. This method has been applied to the ZiYuan-3 satellite system for processing the raw star-tracker and gyro data daily.

Quaternion normalization in additive EKF for spacecraft attitude determination. This work introduces, examines, and compares several quaternion normalization algorithms, which are shown to be an effective stage in the application of the additive extended Kalman filter EKF to spacecraft attitude determination , which is based on vector measurements.

Two new normalization schemes are introduced. They are compared with one another and with the known brute force normalization scheme, and their efficiency is examined. Simulated satellite data are used to demonstrate the performance of all three schemes. A fourth scheme is suggested for future research. Although the schemes were tested for spacecraft attitude determination , the conclusions are general and hold for attitude determination of any three dimensional body when based on vector measurements, and use an additive EKF for estimation, and the quaternion for specifying the attitude.

Orbit determination for ISRO satellite missions. Considering the requirements of satellite missions, software packages are developed, tested and their accuracies are assessed. The results match well with those available from these agencies. These packages have supported orbit determination successfully throughout the mission life for all ISRO satellite missions. Induction magnetometers can further be divided into ei- ther search-coil or fluxgate magnetometers.

A search-coil magnetometer consists of a A fluxgate sensor, which is part of the magnetometer , along with the field control electronics and power supplies compensate for ambient Polat March Polat Due to the mass- and volume-constrained design environment of CubeSat, conventional methods are. This satellite will be designed to fulfil two scientific objectives: The observation of stellar oscillations and the detection and localisation of gamma-ray bursts.

The satellite will be equipped with a tetrahedron Compensation of an attitude disturbance torque caused by magnetic substances in LEO satellites. This research considers an attitude disturbance torque caused by ferromagnetic substances in a LEO satellite. In most LEO satellite missions, a gravity gradient torque, solar pressure torque, aerodynamic torque, and magnetic dipole moment torque are considered for their attitude control systems, however, the effect of the ferromagnetic substances causing a disturbance torque in the geomagnetic field is not considered in previous satellite missions.

The ferromagnetic substances such as iron cores of MTQs and a magnetic hysteresis damper for a passive attitude control system are used in various small satellites. These substances cause a disturbance torque which is almost the same magnitude of the dipole magnetic disturbance and the dominant disturbance in the worst cases. This research proposes a method to estimate and compensate for the effect of the ferromagnetic substances using an extended Kalman filter.

From simulation results, the research concludes that the proposed method is useful and attractive for precise attitude control for LEO satellite missions. Development of a computationally efficient algorithm for attitude estimation of a remote sensing satellite. This paper presents a computationally efficient algorithm for attitude estimation of remote a sensing satellite.

However, utilizing all of the measurement data simultaneously in EKF structure increases computational burden. In order to solve this problem, an efficient version of EKF, namely Murrell's version, is employed. This method utilizes measurements separately at each sampling time for gain computation. Moreover, gyro drifts during the time can reduce the pointing accuracy.

Therefore, a calibration algorithm is utilized for estimation of the main gyro parameters. Full Text Available The paper describes the development of a microsatellite attitude determination and control subsystem ADCS and verification of its functionality by software-in-the-loop SIL method. The role of ADCS is to provide attitude control functions, including the de-tumbling and stabilizing the satellite angular velocity, and as well as estimating the orbit and attitude information during the satellite operation.

During the initialization mode, ADCS collects the early orbit measurement data from various sensors so that the data can be downlinked to the ground station for further analysis. As particularly emphasized in this paper, during the detumbling mode, ADCS implements the thrusters in plus-wide modulation control method to decrease the satellite angular velocity.

ADCS provides the attitude determination function for the estimation of the satellite state, during normal mode. The three modes of microsatellite adopted Kalman filter algorithm estimate microsatellite attitude. Our evaluations are conducted for both a linear array of two antennas and a planar array of three antennas.

A pre-requisite for precise and fast IRNSS attitude determination is the successful resolution of the double-differenced DD integer carrier-phase ambiguities. In this contribution, we will compare the performances of different such methods, amongst which the unconstrained and the multivariate-constrained LAMBDA method for both linear and planar arrays.

Magnetic dipole moment estimation and compensation for an accurate attitude control in nano- satellite missions. Nano- satellites provide space access to broader range of satellite developers and attract interests as an application of the space developments.

These days several new nano- satellite missions are proposed with sophisticated objectives such as remote-sensing and observation of astronomical objects. In these advanced missions, some nano- satellites must meet strict attitude requirements for obtaining scientific data or images. For LEO nano- satellite , a magnetic attitude disturbance dominates over other environmental disturbances as a result of small moment of inertia, and this effect should be cancelled for a precise attitude control.

This research focuses on how to cancel the magnetic disturbance in orbit. This paper presents a unique method to estimate and compensate the residual magnetic moment, which interacts with the geomagnetic field and causes the magnetic disturbance. An extended Kalman filter is used to estimate the magnetic disturbance. This method will be also used for a nano-astrometry satellite mission.

This paper concludes that use of the magnetic disturbance estimation and compensation are useful for nano- satellites missions which require a high accurate attitude control. There is trend in spacecraft engineering toward distributed systems where a number of smaller spacecraft work as a larger satellite. However, in order to make the small satellites work together as a single large platform, the precise relative positions baseline and orientations attitude of the.

The attitude inversion method of geostationary satellites based on unscented particle filter. The attitude information of geostationary satellites is difficult to be obtained since they are presented in non-resolved images on the ground observation equipment in space object surveillance.

In this paper, an attitude inversion method for geostationary satellite based on Unscented Particle Filter UPF and ground photometric data is presented. This update method improves the particle selection based on the idea of UKF to redesign the importance density function. Moreover, it uses the RMS-UKF to partially correct the prediction covariance matrix, which improves the applicability of the attitude inversion method in view of UKF and the particle degradation and dilution of the attitude inversion method based on PF.

This paper describes the main principles and steps of algorithm in detail, correctness, accuracy, stability and applicability of the method are verified by simulation experiment and scaling experiment in the end.

The results show that the proposed method can effectively solve the problem of particle degradation and depletion in the attitude inversion method on account of PF, and the problem that UKF is not suitable for the strong non-linear attitude inversion.

However, the inversion accuracy is obviously superior to UKF and PF, in addition, in the case of the inversion with large attitude error that can inverse the attitude with small particles and high precision. The SMEX program has produced five satellites , three of which have been successfully launched. The remaining two spacecraft are scheduled for launch within the coming year. Next, the context in which the software resides is explained. The paper describes the principles of encapsulation, inheritance, and polymorphism with respect to the implementation of an ACS software system.

This paper will also discuss the design of several ACS software components. Specifically, object-oriented designs are presented for sensor data processing, attitude determination , attitude control, and failure detection. The AFC is a large software repository, requiring a minimal amount of code modifications to produce ACS software for future projects.

This work is based on the previous paper of the author [1]. The present paper is devoted to the investigation of the attitude dynamics of an ecliptic satellite moving in the magnetic field of the Earth. Eelectrodynamic forces result from the motion of a charged satelite relative to the magnetic field of the Earth. The torque due to electrodynamic effect of the Lorentz forces on the attitude stabilization of the satellite is studied with the detailed model of the Earth's magnetic field.

A method for estimating the stable and unstable regions of the equilibrium positions based on Euler's equation is also discussed. The results show that Lorentz forces can affect the stablization of the satellite , in particular for highly eccentric orbits and also for large satellte.

AdSpR 40, , Attractive manifold-based adaptive solar attitude control of satellites in elliptic orbits. The paper presents a novel noncertainty-equivalent adaptive NCEA control system for the pitch attitude control of satellites in elliptic orbits using solar radiation pressure SRP. The satellite is equipped with two identical solar flaps to produce control moments. The control system has a modular controller-estimator structure and has separate tunable gains. A special feature of this NCEA law is that the trajectories of the satellite converge to a manifold in an extended state space, and the adaptive law recovers the performance of a deterministic controller.

This recovery of performance cannot be obtained with certainty-equivalent adaptive CEA laws. Simulation results are presented which show that the NCEA law accomplishes precise attitude control of the satellite in an elliptic orbit, despite large parameter uncertainties. An orbit determination algorithm for small satellites based on the magnitude of the earth magnetic field. Autonomous attitude determination systems based on simple measurements of vector quantities such as magnetic field and the Sun direction are commonly used in very small satellites.

However, those systems always require knowledge of the satellite position. This information can be either propagated from orbital elements periodically uplinked from the ground station or measured onboard by dedicated global positioning system GPS receiver. The former solution sacrifices satellite autonomy while the latter requires additional sensors which may represent a significant part of mass, volume, and power budget in case of pico- or nanosatellites.

Hence, it is thought that a system for onboard satellite position determination without resorting to GPS receivers would be useful. In this paper, a novel algorithm for determining the satellite orbit semimajor-axis is presented. The methods exploit only the magnitude of the Earth magnetic field recorded onboard by magnetometers. This represents the first step toward an extended algorithm that can determine all orbital elements of the satellite.

The method is validated by numerical analysis and real magnetic field measurements. Real time hardware-in-loop simulation of ESMO satellite attitude control system. Two low-power microcontrollers are used for processing and to display information on a graphic screen for real-time performance evaluations.

A new parallel inter-processor communication protocol was developed that is faster and uses less power than existing standard protocols. A shorted annular patch SAP antenna was fabricated for the initial ground-based AD experiments with the testbed. Evaluation of geomagnetic field models using magnetometer measurements for satellite attitude determination system at low earth orbits: Case studies.

In this study, different geomagnetic field models are compared in order to study the errors resulting from the representation of magnetic fields that affect the satellite attitude system. The comparisons were made during geomagnetically active and quiet days to see the effects of the geomagnetic storms and sub-storms on the predicted and observed magnetic fields and angles. The angles, in turn, are used to estimate the spacecraft attitude and hence, the differences between model and observations as well as between two models become important to determine and reduce the errors associated with the models under different space environment conditions.

We show that the models differ from the observations even during the geomagnetically quiet times but the associated errors during the geomagnetically active times increase. We find that the T89 model gives closer predictions to the observations, especially during active times and the errors are smaller compared to the IGRF model. For the first time, the geomagnetic models were used to address the effects of the near Earth space environment on the satellite attitude.

Spacecraft Attitude Determination. This thesis describes the development of an attitude determination system for spacecraft based only on magnetic field measurements. The need for such system is motivated by the increased demands for inexpensive, lightweight solutions for small spacecraft. These spacecraft demands full attitude Meeting these objectives with a single vector magnetometer is difficult and requires temporal fusion of data in order to avoid local observability problems.

In order to guaranteed globally nonsingular solutions, quaternions are generally the preferred attitude The quantitative effects of errors in the process and noise statistics are discussed in detail. Visual attitude propagation for small satellites.

As electronics become smaller and more capable, it has become possible to conduct meaningful and sophisticated satellite missions in a small form factor. However, the capability of small satellites and the range of possible applications are limited by the capabilities of several technologies, including attitude determination and control systems.

This dissertation evaluates the use of image-based visual attitude propagation as a compliment or alternative to other attitude determination technologies that are suitable for miniature satellites. The concept lies in using miniature cameras to track image features across frames and extracting the underlying rotation.

The problem of visual attitude propagation as a small satellite attitude determination system is addressed from several aspects: related work, algorithm design, hardware and performance evaluation, possible applications, and on-orbit experimentation.

These areas of consideration reflect the organization of this dissertation. A "stellar gyroscope" is developed, which is a visual star-based attitude propagator that uses relative motion of stars in an imager's field of view to infer the attitude changes. The device generates spacecraft relative attitude estimates in three degrees of freedom. Algorithms to perform the star detection, correspondence, and attitude propagation are presented. The Random Sample Consensus RANSAC approach is applied to the correspondence problem to successfully pair stars across frames while mitigating falsepositive and false-negative star detections.

This approach provides tolerance to the noise levels expected in using miniature optics and no baffling, and the noise caused by radiation dose on orbit. The hardware design and algorithms are validated using test images of the night sky. The application of the stellar gyroscope as part of a CubeSat attitude determination and control system is described.

The stellar gyroscope is used to augment a MEMS gyroscope attitude propagation. Innovative power management, attitude determination and control tile for CubeSat standard Nano Satellites. Electric power supply EPS and attitude determination and control subsystem ADCS are the most essential elements of any aerospace mission. So keeping in mind their importance, they have been integrated and developed on a single tile called CubePMT module.

Modular power management tiles PMTs are already available in the market but they are less efficient, heavier in weight, consume more power and contain less number of subsystems. CubePMT is developed on the design approach of AraMiS architecture: a project developed at Politecnico di Torino that provides low cost and higher performance space missions with dimensions larger than CubeSats.

The feature of AraMiS design approach is its modularity. These modules can be reused for multiple missions which helps in significant reduction of the overall budget, development and testing time. One has just to reassemble the required subsystems to achieve the targeted specific mission.

Incorporation of star measurements for the determination of orbit and attitude parameters of a geosynchronous satellite : An iterative application of linear regression. The image coordinates of the star locations are measured and stored. Subsequently, the information is used to determine the attitude , the misalignment angles between the spin axis and the principal axis of the satellite , and the precession rate and direction.

This is done for both the 'East' and 'West' operational geosynchronous satellites. This orientation information is then combined with image measurements of earth based landmarks to determine the orbit of each satellite. The method for determining the orbit is simple. For each landmark measurement one determines a nominal position vector for the satellite by extending a ray from the landmark's position towards the satellite and intersecting the ray with a sphere with center coinciding with the Earth's center and with radius equal to the nominal height for a geosynchronous satellite.

The apparent motion of the satellite around the Earth's center is then approximated with a Keplerian model. In turn the variations of the satellite 's height, as a function of time found by using this model, are used to redetermine the successive satellite positions by again using the Earth based landmark measurements and intersecting rays from these landmarks with the newly determined spheres.

This process is performed iteratively until convergence is achieved. Only three iterations are required. Satellite Attitude Control System Simulator. Full Text Available Future space missions will involve satellites with great autonomy and stringent pointing precision, requiring of the Attitude Control Systems ACS with better performance than before, which is function of the control algorithms implemented on board computers.

The difficulties for developing experimental ACS test is to obtain zero gravity and torque free conditions similar to the SCA operate in space. However, prototypes for control algorithms experimental verification are fundamental for space mission success.

This paper presents the parameters estimation such as inertia matrix and position of mass centre of a Satellite Attitude Control System Simulator SACSS, using algorithms based on least square regression and least square recursive methods. Simulations have shown that both methods have estimated the system parameters with small error.

The SACSS platform model will be used to do experimental verification of fundamental aspects of the satellite attitude dynamics and design of different attitude control algorithm. As three axis stabilization is high level technology requiring the proper functioning of various sensors, actuators and control software, many early satellites failed in their initial operation phase because of shortage of solar power generation or inability to realize the initial step of missions because of unexpected attitude control system performance.

These results come from failure to design the satellite attitude determination and control system ADCS appropriately and not considering " satellite survivability. This paper discusses how to realize ADCS while taking satellite survivability into account, based on our experiences of design and in-orbit operations of Hodoyoshi-3 and 4 satellites launched in , which suffered from various component anomalies but could complete their missions.

Passive Magnetic Attitude Control PMAC is capable of aligning a satellite within 5 degrees of the local magnetic field at low resource cost, making it ideal for a small satellite. However, simulation attempts to date have not been able to predict the attitude dynamics at a level sufficient for mission design. Also, some satellites have suffered from degraded performance due to an incomplete understanding of PMAC system design.

This dissertation alleviates these issues by discussing the design, inputs, and validation of PMAC systems for small satellites. After on-orbit calibration of the off-the-shelf magnetometer and photodiodes and an on-orbit fit to the satellite magnetic moment, the MEKF regularly achieves a three sigma attitude uncertainty of 4 degrees or less. CSSWE is found to settle to the magnetic field in seven days, verifying its attitude design requirement. A Helmholtz cage is constructed and used to characterize the CSSWE bar magnet and hysteresis rods both individually and in the flight configuration.

Fitted parameters which govern the magnetic material behavior are used as input to a PMAC dynamics simulation. All components of this simulation are described and defined. Simulation-based dynamics analysis shows that certain initial conditions result in abnormally decreased settling times; these cases may be identified by their dynamic response.

The simulation output is compared to the MEKF output; the true dynamics are well modeled and the predicted settling time is found to possess a 20 percent error, a significant improvement over prior simulation. Satellite recovery - Attitude dynamics of the targets. The problems of categorizing and modeling the attitude dynamics of uncontrolled artificial earth satellites which may be targets in recovery attempts are addressed. Methods of classification presented are based on satellite rotational kinetic energy, rotational angular momentum and orbit and on the type of control present prior to the benign failure of the control system.

The use of approximate analytical solutions and 'exact' numerical solutions to the equations governing satellite attitude motions to predict uncontrolled attitude motion is considered. Analytical and numerical results are presented for the evolution of satellite attitude motions after active control termination. While there are a multitude of ways to determine a satellite 's orientation, very little research has been done on determining if the attitude of a satellite can be determined directly from telemetry The three ACTS control axes are defined, including the means for sensing attitude and determining the pointing errors.

The desired pointing requirements for various modes of control as well as the disturbance torques that oppose the control are identified. Finally, the hardware actuators and control loops utilized to reduce the attitude error are described.

A new star tracker concept for satellite attitude determination based on a multi-purpose panoramic camera. This paper presents an innovative algorithm developed for attitude determination of a space platform. The algorithm exploits images taken from a multi-purpose panoramic camera equipped with hyper-hemispheric lens and used as star tracker.

The sensor architecture is also original since state-of-the-art star trackers accurately image as many stars as possible within a narrow- or medium-size field-of-view, while the considered sensor observes an extremely large portion of the celestial sphere but its observation capabilities are limited by the features of the optical system. The proposed original approach combines algorithmic concepts, like template matching and point cloud registration, inherited from the computer vision and robotic research fields, to carry out star identification.

The final aim is to provide a robust and reliable initial attitude solution lost-in-space mode , with a satisfactory accuracy level in view of the multi-purpose functionality of the sensor and considering its limitations in terms of resolution and sensitivity. Performance evaluation is carried out within a simulation environment in which the panoramic camera operation is realistically reproduced, including perturbations in the imaged star pattern.

Confined computer capacity and a limit on electrical power supply were separate obstacles. They demanded computational simplicity and power optimality from the attitude control system. The design of quasi optimal controllers for a real-time implementation It was explained how the equilibria depended on the ratio of the satellite 's moments of inertia. It was further investigated how to control the attitude , such that the satellite was globally asymptotically stable in the desired Satellite Photometric Error Determination.

Payne, Philip J. Castro, Stephen A. The Johnson photometric system is a set of filters in the optical. Attitude stability analyses for small artificial satellites. The objective of this paper is to analyze the stability of the rotational motion of a symmetrical spacecraft, in a circular orbit.

The equilibrium points and regions of stability are established when components of the gravity gradient torque acting on the spacecraft are included in the equations of rotational motion, which are described by the Andoyer's variables.

The nonlinear stability of the equilibrium points of the rotational motion is analysed here by the Kovalev-Savchenko theorem. With the application of the Kovalev-Savchenko theorem, it is possible to verify if they remain stable under the influence of the terms of higher order of the normal Hamiltonian. In this paper, numerical simulations are made for a small hypothetical artificial satellite.

Several stable equilibrium points were determined and regions around these points have been established by variations in the orbital inclination and in the spacecraft principal moment of inertia. The present analysis can directly contribute in the maintenance of the spacecraft's attitude.

Statistical Attitude Determination. All spacecraft require attitude determination at some level of accuracy. This can be a very coarse requirement of tens of degrees, in order to point solar arrays at the sun, or a very fine requirement in the milliarcsecond range, as required by Hubble Space Telescope. A toolbox of attitude determination methods, applicable across this wide range, has been developed over the years. There have been many advances in the thirty years since the publication of Reference, but the fundamentals remain the same.

One significant change is that onboard attitude determination has largely superseded ground-based attitude determination , due to the greatly increased power of onboard computers. The availability of relatively inexpensive radiation-hardened microprocessors has led to the development of "smart" sensors, with autonomous star trackers being the first spacecraft application.

Another new development is attitude determination using interferometry of radio signals from the Global Positioning System GPS constellation. This article reviews both the classic material and these newer developments at approximately the level of, with emphasis on. We discuss both "single frame" methods that are based on measurements taken at a single point in time, and sequential methods that use information about spacecraft dynamics to combine the information from a time series of measurements.

Introducing the state variable function of GPS attitude determination algorithm in SAR satellite by means of kinematic vector and describing the observation function by the GPS wide-band carrier phase, the paper uses the Kalman filter algorithm to obtian the attitude variables of SAR satellite. Compared the simulation results of Kalman filter algorithm with the least square algorithm and explicit solution, it is indicated that the Kalman filter algorithm is the best. Full Text Available This article has discussed the development of a three-axis attitude digital controller for an artificial satellite using a digital signal processor.

The main motivation of this study is the attitude control system of the satellite Multi-Mission Platform, developed by the Brazilian National Institute for Space Research for application in different sort of missions. The controller design was based on the theory of the Linear Quadratic Gaussian Regulator, synthesized from the linearized model of the motion of the satellite , i.

In the first stage of the project development, a system controller for continuous time was studied with the aim of testing the adequacy of the adopted control. Star trackers for attitude determination. One problem comes to all spacecrafts using vector information. That is the problem of determining the attitude.

This paper describes how the area of attitude determination instruments has evolved from simple pointing devices into the latest technology, which determines the attitude by utilizing The instruments are called star trackers and they are capable of determining the attitude with an accuracy better than 1 arcsecond. The concept of the star tracker is explained. The obtainable accuracy is calculated, the numbers of stars to be included Finally the commercial market for star trackers is discussed Touchless attitude correction for satellite with constant magnetic moment.

Rescue of satellite with attitude fault is of great value. Satellite with improper injection attitude may lose contact with ground as the antenna points to the wrong direction, or encounter energy problems as solar arrays are not facing the sun. Improper uploaded command may set the attitude out of control, exemplified by Japanese Hitomi spacecraft.

In engineering practice, traditional physical contact approaches have been applied, yet with a potential risk of collision and a lack of versatility since the mechanical systems are mission-specific. This paper puts forward a touchless attitude correction approach, in which three satellites are considered, one having constant dipole and two having magnetic coils to control attitude of the first.

Particular correction configurations are designed and analyzed to maintain the target's orbit during the attitude correction process. A reference coordinate system is introduced to simplify the control process and avoid the singular value problem of Euler angles.

Based on the spherical triangle basic relations, the accurate varying geomagnetic field is considered in the attitude dynamic mode. Sliding mode control method is utilized to design the correction law. Finally, numerical simulation is conducted to verify the theoretical derivation.

It can be safely concluded that the no-contact attitude correction approach for the satellite with uniaxial constant magnetic moment is feasible and potentially applicable to on-orbit operations. Integrated orbit and attitude hardware-in-the-loop simulations for autonomous satellite formation flying. Development and experiment of an integrated orbit and attitude hardware-in-the-loop HIL simulator for autonomous satellite formation flying are presented.

The integrated simulator system consists of an orbit HIL simulator for orbit determination and control, and an attitude HIL simulator for attitude determination and control. The integrated simulator involves four processes orbit determination , orbit control, attitude determination , and attitude control , which interact with each other in the same way as actual flight processes do. Orbit determination is conducted by a relative navigation algorithm using double-difference GPS measurements based on the extended Kalman filter EKF.

Orbit control is performed by a state-dependent Riccati equation SDRE technique that is utilized as a nonlinear controller for the formation control problem. Attitude is determined from an attitude heading reference system AHRS sensor, and a proportional-derivative PD feedback controller is used to control the attitude HIL simulator using three momentum wheel assemblies. Integrated orbit and attitude simulations are performed for a formation reconfiguration scenario.

By performing the four processes adequately, the desired formation reconfiguration from a baseline of m was achieved with meter-level position error and millimeter-level relative position navigation. This HIL simulation demonstrates the performance of the integrated HIL simulator and the feasibility of the applied algorithms in a real-time environment. Furthermore, the integrated HIL simulator system developed in the current study can be used as a ground-based testing environment to reproduce possible actual satellite formation operations.

Sensor fault detection and recovery in satellite attitude control. This paper proposes an integrated sensor fault detection and recovery for the satellite attitude control system. By introducing a nonlinear observer, the healthy sensor measurements are provided. Considering attitude dynamics and kinematic, a novel observer is developed to detect the fault in angular rate as well as attitude sensors individually or simultaneously.

There is no limit on type and configuration of attitude sensors. By designing a state feedback based control signal and Lyapunov stability criterion, the uniformly ultimately boundedness of tracking errors in the presence of sensor faults is guaranteed. Finally, simulation results are presented to illustrate the performance of the integrated scheme. This has been a serious obstacle for using magnetorquer based control for three-axis attitude control.

This paper deals with three-axis stabilization of a low earth orbit satellite. The problem of controlling A three dimensional sliding manifold is proposed, and it is shown that the satellite motion on the sliding manifold is asymptotically stable Low-power attitude determination for magnetometry planetary missions.

This work covers the subject of orientation or attitude in space and on the surface of a planet. Different attitude sensor technologies have been investigated with emphasis on very low power consumption and mass. In addition robust methods for attitude determination have been covered again A true low-power attitude sensor using the Anisotropic Magneto Resistor effect have been designed to late prototype state.

Two prototypes of the AMR magnetometer have been built. One of the prototypes has an analog output and the second Different attitude representations such as orthogonal matrices, Euler angles and quaternions are presented.

Also methods for attitude determination of a sensor platform with more than one vector Performance comparison of attitude determination , attitude estimation, and nonlinear observers algorithms. This paper presents a brief synthesis and useful performance analysis of different attitude filtering algorithms attitude determination algorithms, attitude estimation algorithms, and nonlinear observers applied to Low Earth Orbit Satellite in terms of accuracy, convergence time, amount of memory, and computation time.

This latter is calculated in two ways, using a personal computer and also using On-board computer OBC that is being used in many SSTL Earth observation missions. The use of this comparative study could be an aided design tool to the designer to choose from an attitude determination or attitude estimation or attitude observer algorithms. The simulation results clearly indicate that the nonlinear Observer is the more logical choice.

Hughey, Raymond H. Cost, size, weight and power requirements, on the other hand, impose selecting relative simple sensors and actuators which leads to an attitude control requirement of less than 1 degree. This precision is obtained by a combination of magnetorquers and momentum wheels Attitude Determination and Control Systems. In the year , Galveston, Texas, was a bustling community of approximately 40, people.

The former capital of the Republic of Texas remained a trade center for the state and was one of the largest cotton ports in the United States. On September 8 of that year, however, a powerful hurricane struck Galveston island, tearing the Weather Bureau wind gauge away as the winds exceeded mph and bringing a storm surge that flooded the entire city. The worst natural disaster in United States history even today the hurricane caused the deaths of between and people. Critical in the events that led to such a terrible loss of life was the lack of precise knowledge of the strength of the storm before it hit.

In , Hurricane Ike, the third costliest hurricane ever to hit the United States coast, traveled through the Gulf of Mexico. Ike was gigantic, and the devastation in its path included the Turk and Caicos Islands, Haiti, and huge swaths of the coast of the Gulf of Mexico. Once again, Galveston, now a city of nearly 60,, took the direct hit as Ike came ashore.

Almost people in the Caribbean and the United States lost their lives; a tragedy to be sure, but far less deadly than the storm. This time, people were prepared, having received excellent warning from the GOES satellite network. The Geostationary Operational Environmental Satellites have been a continuous monitor of the world's weather since , and they have since been joined by other Earth-observing satellites.

This weather surveillance to which so many now owe their lives is possible in part because of the ability to point accurately and steadily at the Earth below. The importance of accurately pointing spacecraft to our daily lives is pervasive, yet somehow escapes the notice of most people.

But the example of the lives saved from Hurricane Ike as compared to the storm is something no one should ignore. In this section, we will summarize the processes and technologies used in. Development of a hardware-in-loop attitude control simulator for a CubeSat satellite. Attitude control is an important part in satellite on-orbit operation.

It greatly affects the performance of satellites. Testing of an attitude determination and control subsystem ADCS is very challenging since it might require attitude dynamics and space environment in the orbit. The simulator consists of a numerical simulation part, a hardware part, and a HIL interface hardware unit.

The hardware part is the real ADCS board of the satellite. Then, based on this information, the HIL interface hardware generates I2C signals mimicking the signals of the on-board rate-gyros and magnetometers and consequently outputs the signals to the ADCS board. The ADCS board reads the rate-gyro and magnetometer signals, calculates control signals, and drives the attitude actuators which are three magnetic torquers MTQs.

The responses of the MTQs sensed by a separated magnetometer are feedback to the numerical simulation part completing the HIL simulation loop. Experimental studies are conducted to demonstrate the feasibility and effectiveness of the simulator. Full Text Available Attitude data, which is the important data strongly correlated with the geometric accuracy of optical remote sensing satellite images, are generally obtained using a real-time Extended Kalman Filter EKF with star-tracker and gyro data for current high-resolution satellites , such as Orb-view, IKONOS, Quickbird,Pleiades, and ZY Moreover, this method makes full use of the collected data in the fixed-interval and computational resources on the ground, and it determines optimal attitude results by forward-backward filtering and weighted smoothing with the raw star-tracker and gyro data collected for a fixed period.

In this study, the principle and implementation of the proposed method are described. The post-processed attitude was compared with the on-board attitude , and the absolute accuracy was evaluated by the two methods.

One method compares the positioning accuracy of the object space coordinates with the post-processed and on-board attitude data without using ground control points GCPs. The other method compares the tie-point residuals of the image coordinates after a free net adjustment. In addition, the internal and external parameters of the camera were accurately calibrated before use for an objective evaluation of the attitude accuracy.

The experimental results reveal that the accuracy of the post-processed attitude is superior to the accuracy of the on-board processed attitude. This method has been applied to the ZiYuan-3 satellite system for processing the raw star-tracker and gyro data daily. Quaternion normalization in additive EKF for spacecraft attitude determination. This work introduces, examines, and compares several quaternion normalization algorithms, which are shown to be an effective stage in the application of the additive extended Kalman filter EKF to spacecraft attitude determination , which is based on vector measurements.

Two new normalization schemes are introduced. They are compared with one another and with the known brute force normalization scheme, and their efficiency is examined. Simulated satellite data are used to demonstrate the performance of all three schemes. A fourth scheme is suggested for future research. Although the schemes were tested for spacecraft attitude determination , the conclusions are general and hold for attitude determination of any three dimensional body when based on vector measurements, and use an additive EKF for estimation, and the quaternion for specifying the attitude.

Orbit determination for ISRO satellite missions. Considering the requirements of satellite missions, software packages are developed, tested and their accuracies are assessed. The results match well with those available from these agencies.

These packages have supported orbit determination successfully throughout the mission life for all ISRO satellite missions. Induction magnetometers can further be divided into ei- ther search-coil or fluxgate magnetometers. A search-coil magnetometer consists of a A fluxgate sensor, which is part of the magnetometer , along with the field control electronics and power supplies compensate for ambient Polat March Polat Due to the mass- and volume-constrained design environment of CubeSat, conventional methods are.

This satellite will be designed to fulfil two scientific objectives: The observation of stellar oscillations and the detection and localisation of gamma-ray bursts. The satellite will be equipped with a tetrahedron Compensation of an attitude disturbance torque caused by magnetic substances in LEO satellites.

This research considers an attitude disturbance torque caused by ferromagnetic substances in a LEO satellite. In most LEO satellite missions, a gravity gradient torque, solar pressure torque, aerodynamic torque, and magnetic dipole moment torque are considered for their attitude control systems, however, the effect of the ferromagnetic substances causing a disturbance torque in the geomagnetic field is not considered in previous satellite missions.

The ferromagnetic substances such as iron cores of MTQs and a magnetic hysteresis damper for a passive attitude control system are used in various small satellites. These substances cause a disturbance torque which is almost the same magnitude of the dipole magnetic disturbance and the dominant disturbance in the worst cases.

This research proposes a method to estimate and compensate for the effect of the ferromagnetic substances using an extended Kalman filter. From simulation results, the research concludes that the proposed method is useful and attractive for precise attitude control for LEO satellite missions.

Development of a computationally efficient algorithm for attitude estimation of a remote sensing satellite. This paper presents a computationally efficient algorithm for attitude estimation of remote a sensing satellite. However, utilizing all of the measurement data simultaneously in EKF structure increases computational burden.

In order to solve this problem, an efficient version of EKF, namely Murrell's version, is employed. This method utilizes measurements separately at each sampling time for gain computation. Moreover, gyro drifts during the time can reduce the pointing accuracy. Therefore, a calibration algorithm is utilized for estimation of the main gyro parameters. Full Text Available The paper describes the development of a microsatellite attitude determination and control subsystem ADCS and verification of its functionality by software-in-the-loop SIL method.

The role of ADCS is to provide attitude control functions, including the de-tumbling and stabilizing the satellite angular velocity, and as well as estimating the orbit and attitude information during the satellite operation. During the initialization mode, ADCS collects the early orbit measurement data from various sensors so that the data can be downlinked to the ground station for further analysis. As particularly emphasized in this paper, during the detumbling mode, ADCS implements the thrusters in plus-wide modulation control method to decrease the satellite angular velocity.

ADCS provides the attitude determination function for the estimation of the satellite state, during normal mode. The three modes of microsatellite adopted Kalman filter algorithm estimate microsatellite attitude. Our evaluations are conducted for both a linear array of two antennas and a planar array of three antennas. A pre-requisite for precise and fast IRNSS attitude determination is the successful resolution of the double-differenced DD integer carrier-phase ambiguities.

In this contribution, we will compare the performances of different such methods, amongst which the unconstrained and the multivariate-constrained LAMBDA method for both linear and planar arrays. Magnetic dipole moment estimation and compensation for an accurate attitude control in nano- satellite missions.

Nano- satellites provide space access to broader range of satellite developers and attract interests as an application of the space developments. These days several new nano- satellite missions are proposed with sophisticated objectives such as remote-sensing and observation of astronomical objects. In these advanced missions, some nano- satellites must meet strict attitude requirements for obtaining scientific data or images.

For LEO nano- satellite , a magnetic attitude disturbance dominates over other environmental disturbances as a result of small moment of inertia, and this effect should be cancelled for a precise attitude control. This research focuses on how to cancel the magnetic disturbance in orbit. This paper presents a unique method to estimate and compensate the residual magnetic moment, which interacts with the geomagnetic field and causes the magnetic disturbance.

An extended Kalman filter is used to estimate the magnetic disturbance. This method will be also used for a nano-astrometry satellite mission. This paper concludes that use of the magnetic disturbance estimation and compensation are useful for nano- satellites missions which require a high accurate attitude control.

There is trend in spacecraft engineering toward distributed systems where a number of smaller spacecraft work as a larger satellite. However, in order to make the small satellites work together as a single large platform, the precise relative positions baseline and orientations attitude of the.

The attitude inversion method of geostationary satellites based on unscented particle filter. The attitude information of geostationary satellites is difficult to be obtained since they are presented in non-resolved images on the ground observation equipment in space object surveillance.

In this paper, an attitude inversion method for geostationary satellite based on Unscented Particle Filter UPF and ground photometric data is presented. This update method improves the particle selection based on the idea of UKF to redesign the importance density function.

Moreover, it uses the RMS-UKF to partially correct the prediction covariance matrix, which improves the applicability of the attitude inversion method in view of UKF and the particle degradation and dilution of the attitude inversion method based on PF. This paper describes the main principles and steps of algorithm in detail, correctness, accuracy, stability and applicability of the method are verified by simulation experiment and scaling experiment in the end. The results show that the proposed method can effectively solve the problem of particle degradation and depletion in the attitude inversion method on account of PF, and the problem that UKF is not suitable for the strong non-linear attitude inversion.

However, the inversion accuracy is obviously superior to UKF and PF, in addition, in the case of the inversion with large attitude error that can inverse the attitude with small particles and high precision. The SMEX program has produced five satellites , three of which have been successfully launched. The remaining two spacecraft are scheduled for launch within the coming year.

Next, the context in which the software resides is explained. The paper describes the principles of encapsulation, inheritance, and polymorphism with respect to the implementation of an ACS software system. This paper will also discuss the design of several ACS software components. Specifically, object-oriented designs are presented for sensor data processing, attitude determination , attitude control, and failure detection.

The AFC is a large software repository, requiring a minimal amount of code modifications to produce ACS software for future projects. This work is based on the previous paper of the author [1]. The present paper is devoted to the investigation of the attitude dynamics of an ecliptic satellite moving in the magnetic field of the Earth.

Eelectrodynamic forces result from the motion of a charged satelite relative to the magnetic field of the Earth. The torque due to electrodynamic effect of the Lorentz forces on the attitude stabilization of the satellite is studied with the detailed model of the Earth's magnetic field. A method for estimating the stable and unstable regions of the equilibrium positions based on Euler's equation is also discussed.

The results show that Lorentz forces can affect the stablization of the satellite , in particular for highly eccentric orbits and also for large satellte. AdSpR 40, , Attractive manifold-based adaptive solar attitude control of satellites in elliptic orbits. The paper presents a novel noncertainty-equivalent adaptive NCEA control system for the pitch attitude control of satellites in elliptic orbits using solar radiation pressure SRP.

The satellite is equipped with two identical solar flaps to produce control moments. The control system has a modular controller-estimator structure and has separate tunable gains. A special feature of this NCEA law is that the trajectories of the satellite converge to a manifold in an extended state space, and the adaptive law recovers the performance of a deterministic controller. This recovery of performance cannot be obtained with certainty-equivalent adaptive CEA laws.

Simulation results are presented which show that the NCEA law accomplishes precise attitude control of the satellite in an elliptic orbit, despite large parameter uncertainties. An orbit determination algorithm for small satellites based on the magnitude of the earth magnetic field.

Autonomous attitude determination systems based on simple measurements of vector quantities such as magnetic field and the Sun direction are commonly used in very small satellites. However, those systems always require knowledge of the satellite position. This information can be either propagated from orbital elements periodically uplinked from the ground station or measured onboard by dedicated global positioning system GPS receiver.

The former solution sacrifices satellite autonomy while the latter requires additional sensors which may represent a significant part of mass, volume, and power budget in case of pico- or nanosatellites. Hence, it is thought that a system for onboard satellite position determination without resorting to GPS receivers would be useful.

In this paper, a novel algorithm for determining the satellite orbit semimajor-axis is presented. The methods exploit only the magnitude of the Earth magnetic field recorded onboard by magnetometers. This represents the first step toward an extended algorithm that can determine all orbital elements of the satellite.

The method is validated by numerical analysis and real magnetic field measurements. Real time hardware-in-loop simulation of ESMO satellite attitude control system. Bang-bang control with dead-zone and Pulse-Width Modulation PWM for the modulation of the on-time of the thrusters are treated. Results for real time simulation are compared with autonomous simulations. The controller gives a satisfactory performance in the real time environment.

This paper discusser linear attitude control strategies for a low earth orbit satellite actuated by a set of mutually perpendicular electromagnetic coils. The principle is to use the interaction between the Earth's magnetic field An observation that geomagnetic field changes approximately periodically when satellite is on a near polar orbit is used throughout this paper. Three types of attitude controllers are proposed Their performance is evaluated and compared in the simulation study of the environment National Aeronautics and Space Administration — To determine pointing and position vectors in both local and inertial coordinate frames, multi-spacecraft missions typically utilize separate attitude determination The determination of the attitude and attitude dynamics of TeamSat.

AVS successfully Full Text Available In this paper, an adaptive attitude control algorithm is developed based on neural network for a satellite using four reaction wheels in a tetrahedron configuration. Then, an attitude control based on feedback linearization control has been designed and uncertainties in the moment of inertia matrix and disturbances torque have been considered.

In order to eliminate the effect of these uncertainties, a multilayer neural network with back-propagation law is designed. In this structure, the parameters of the moment of inertia matrix and external disturbances are estimated and used in feedback linearization control law. Finally, the performance of the designed attitude controller is investigated by several simulations. One of the most challenging problems for marine satellite tracking antennas MSTAs is to estimate the antenna attitude , which is affected by the ship motion, especially the ship vibration and rotational motions caused by ocean waves.

To overcome this problem, an attitude heading and reference We present a study on noise in space-based photometry originating from sensitivity variations within individual pixels, known as intra-pixel variations, and satellite attitude jitter. We have measured the intra-pixel structure on an e2v CCD and made simulations of the effects these structur The purpose of this study is to explore teacher attitudes toward gifted students in several distinct areas and to provide psychometric evidence of reliability and validity for the use of an instrument titled " Determining Attitudes Toward Ability" DATA to measure specific components of teacher attitudes.

Subscales of Focus on Others,…. Performance analysis of a GPS Interferometric attitude determination system for a gravity gradient stabilized spacecraft. The performance of an unaided attitude determination system based on GPS interferometry is examined using linear covariance analysis. The principal error sources are identified and modelled. The optimal system's sensitivities to these error sources are examined through an error budget and by varying system parameters.

The attitude performance of two optimal-suboptimal filters is also presented. Based on this analysis, the limiting factors in attitude accuracy are the knowledge of the relative antenna locations, the electrical path lengths from the antennae to the receiver, and the multipath environment. The performance of the system is found to be fairly insensitive to torque errors, orbital inclination, and the two satellite geometry figures-of-merit tested.

GNSS satellite transmit power and its impact on orbit determination. Antenna thrust is a small acceleration acting on Global Navigation Satellite System satellites caused by the transmission of radio navigation signals. Knowledge about the transmit power and the mass of the satellites is required for the computation of this effect.

The actual transmit power can be obtained from measurements with a high-gain antenna and knowledge about the properties of the transmit and receive antennas as well as losses along the propagation path. The transmit power differs usually only slightly for individual spacecraft within one satellite block. Considering the antenna thrust in precise orbit determination of GNSS satellites decreases the orbital radius by mm depending on the transmit power, the satellite mass, and the orbital period.

Full Text Available In this paper a controllability study of different actuator configurations consisting of magnetic torquers, reaction wheels and a gravity boom is presented. The theoretical analysis is performed with use of controllability gramians, and simulation results with the different configurations are presented and compared regarding settling time and power consumption to substantiate the theoretical analysis.

A reference model is also introduced to show how the power consumption can he lowered to the same magnitude as when magnetic torquers are used, without degrading the satellite response significantly. Global navigation satellite systems GNSS are well suited for attitude determination. In this study, we use the rotation matrix method to resolve the attitude angle. This method achieves better performance in reducing computational complexity and selecting satellites. The condition of the baseline length is combined with the ambiguity function method AFM to search for integer ambiguity, and it is validated in reducing the span of candidates.

The noise error is always the key factor to the success rate. It is closely related to the satellite geometry model. Although the AFM is more flexible, it is lack of analysis on this aspect. In this study, the influence of the satellite geometry model on the success rate is analyzed in detail. The computation error and the noise error are effectively treated. Not only is the flexibility of the AFM inherited, but the success rate is also increased.

An experiment is conducted in a selected campus, and the performance is proved to be effective. Our results are based on simulated and real-time GNSS data and are applied on single-frequency processing, which is known as one of the challenging case of GNSS attitude determination. By employing passivity theory it is shown, that the satellite is a passive system.

This paper shows, that global asymptotic can be obtained with a passive and an imput It is demonstrated in a simulation study that the resultant control has a potential for on-board implementation in the acquistion phase, where global stabillity of the control law is vital Full Text Available In order to realize the high accuracy attitude control of satellite with flexible appendages, attitude control system consisting of the controller and structural filter was designed.

When the low order vibration frequency of flexible appendages is approximating the bandwidth of attitude control system, the vibration signal will enter the control system through measurement device to bring impact on the accuracy or even the stability. In order to reduce the impact of vibration of appendages on the attitude control system, the structural filter is designed in terms of rejecting the vibration of flexible appendages.

Considering the potential problem of in-orbit frequency variation of the flexible appendages, the design method for the adaptive notch filter is proposed based on the in-orbit identification technology. Finally, the simulation results are given to demonstrate the feasibility and effectiveness of the proposed design techniques.

The mathematical specifications of Release 4.

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