Spacecraft Attitude Determination System Research Articles

Spacecraft attitude testbed

Spacecraft attitude determination and control systems (ADCS) stand as a pivotal factor for ensuring the success of missions. As satellite sizes continue to decrease, the need for more intelligent and efficient control algorithms becomes increasingly pronounced. In response to this imperative, we propose the development of a scaled-down testbed specifically designed for evaluating satellites (ACDS). This testbed will demonstrate effective control of a CubeSat through the use of Helmholtz coils, magnetic torquer solenoids, reaction wheels and inexpensive student-made components and available components off the shelf (COTS) in the San Diego State University SPACE Lab. Referred to as the Spacecraft Attitude Testbed (SAT), this custom-designed 3-DoF experimental attitude platform replicates the conditions experienced by a weightless satellite in space. The objective of this research paper is to improve the understanding of how to manufacture a testbed and components like those within this paper, expose some of the issues encountered and improve the outcomes for future academic testbeds. Within this study, the testbed employs torque application to magnetic torquers and reaction wheels to regulate attitude. In summary, this research endeavors to establish a comprehensive and cost-effective testbed platform dedicated to the assessment of spacecraft attitude control and determination systems. It not only seeks to enhance our understanding of advanced control algorithms but also aims to contribute valuable insights into fault detection and adaptive control strategies, ultimately bolstering the efficiency and robustness of satellite operations.

Sun-pointing safe control for atmospheric environment monitoring satellite using magnetic actuation

The design of sun-pointing safe controller is an important part of spacecraft attitude determination and control system. In this work, a new spin stabilization algorithm using only magnetic torquers is proposed. This algorithm has been developed for the atmospheric environment monitoring satellite, which completes the missions of fixed axis rotation and tracking the sun orientation. Firstly, a Lyapunov function, combining the angular momentum error and angular velocity, is formulated and a new sun-pointing attitude control scheme is derived. Then, it is proved that the magnetic control law is asymptotically stable in the absence of disturbance torques. The proposed method is compared with a state-of-the-art fault-tolerant magnetic spin stabilizing controller design method by numerical simulations. Simulation results show that the proposed method has better performance than the state-of-the-art technique. Finally, the applicability and effectiveness of the proposed control method are demonstrated by the hardware-in-the-loop simulation of the satellite.

Centered Error Entropy-Based Sigma-Point Kalman Filter for Spacecraft State Estimation with Non-Gaussian Noise

The classical sigma-point Kalman filter (SPKF) is widely used in a spacecraft state estimation area with the Gaussian white noise hypothesis. The actual sensor noise is often disturbed by outliers in the harsh space environment, and the SPKF algorithm will reduce the filtering accuracy or even diverge. In this study, to enhance the robustness under non-Gaussian noise condition, the outlier-robust SPKF algorithm based on a centered error entropy (CEE) criterion is derived. Unscented Kalman filter (UKF) is typical of SPKF; combining the deterministic sampling criterion with the centered error entropy criterion, a robust centered error entropy UKF (CEEUKF) algorithm is proposed. The CEEUKF uses the unscented transformation (UT) method to perform time update step and then uses the robust regression model and CEE criterion to reconstruct the measurement update step. The effectiveness of the proposed CEEUKF is verified by a spacecraft attitude determination system.

Integrated design and accuracy analysis of star sensor and gyro on the same benchmark for satellite attitude determination system

As an important sensor in the navigation systems, star sensors and the gyro play important roles in spacecraft attitude determination system. Complex environmental factors are the main sources of error in attitude determination. The error influence of different benchmarks and the disintegration mode between the star sensor and the gyro is analyzed in theory. The integrated design of the star sensor and the gyro on the same benchmark can effectively avoid the error influence and improves the spacecraft attitude determination accuracy. Simulation results indicate that when the stars sensor optical axis vectors overlap the reference coordinate axis of the gyro in the same benchmark, the attitude determination accuracy improves.

Multiple-model adaptive estimator for spacecraft attitude sensor calibration

PurposeThis paper aims to present a multiple-model adaptive estimator (MMAE) to calibrate the star sensor low frequency error (LFE). The star sensor LFE, which is caused primarily by the periodic thermal distortion, has a great impact on spacecraft attitude determination accuracy.Design/methodology/approachThe unfavorable effect of the LFE can be partly eliminated by using the calibration algorithm based on the augmented Kalman filter (AKF). However, the AKF may be worse than the traditional Kalman filter (KF) in the absence of the LFE. To cope with this problem, the MMAE is applied first time for combining the AKF and the KF in the spacecraft attitude determination system, such that satisfactory performance can be achieved in different operating scenarios.FindingsThe convergence of the presented MMAE is demonstrated through a formal derivation. A novel method is proposed to tune the MMAE design parameter, such that the convergence rate of the estimator is increased. It is shown via numerical studies that the presented algorithm outperforms the AKF and the KF.Practical implicationsThe calibration algorithm is applicable for spacecraft attitude determination.Originality/valueAn effective star sensor LFE calibration algorithm based on the MMAE is developed. In addition, a novel method is proposed to increase convergence rate of the estimator.

Supervised learning algorithms for spacecraft attitude determination and control system health monitoring

The main concern of any space mission operation is to ensure the health and safety of spacecrafts. The worst case in this circumstance is probably the loss of a mission but the more common interruption of spacecraft functionality can result in compromised mission objectives. Spacecraft telemetry holds information related to state-of-health (SOH) of its subsystems. Each parameter has information that represents a time-variant property (i.e., a status or a measurement) to be checked. Moreover, the tremendous increase in telemetry data volume and its complexity directs the need for more efficient and scalable data processing systems. As a consequence, there is a continuous improvement of telemetry monitoring applications in order to reduce the time required to respond to changes in a spacecraft's SOH. So, a fast conception of the current state of the spacecraft is thus very important in order to respond to failures. Furthermore, the progressive growth in spacecrafts leads to increase of the level of standardization in spacecraft operations.

Robust H∞ filtering for a spacecraft attitude determination system with affine LPV approach

A novel approach is presented in this paper for the filtering design in a spacecraft attitude determination system (ADS), which is equipped with a star sensor and three-axis gyros. Firstly, the ADS workflow and its mathematical model components are described. Then, the ADS state-space model is established as a linear parameter varying (LPV) system, and its matrices are assumed to affinely depend on time-varying parameters, i.e. three-axis angular rates. Secondly, system uncertainties and non-Gaussian noises in the spacecraft ADS are considered, and a parameter-dependent robust H∞ filter is designed based on the established LPV model. This filtering design is then transformed as an optimisation process in terms of linear matrix inequalities (LMIs) which is numerically tractable. At last, results from both simulations and hardware-in-the-loop tests with simulated star-tracker measurements are presented. It can be observed that, compared with the Kalman filter, the proposed filtering method achieves more robust results with considering model uncertainties and non-Gaussian noises, which is closer to practical situations. Also, it is more computationally efficient than the Kalman filter, which is very demanding with hard real-time requirement.

Decentralized diagnosis in a spacecraft attitude determination and control system

In model-based diagnosis (MBD), structural models can provide useful information for fault diagnosis and fault-tolerant control design. In particular, they are known for supporting the design of analytical redundancy relations (ARRs) which are widely used to generate residuals for diagnosis. On the other hand, systems are increasingly complex whereby it is necessary to develop decentralized architectures to perform the diagnosis task. Decentralized diagnosis is of interest for on-board systems as a way to reduce computational costs or for large geographically distributed systems that require to minimizing data transfer. Decentralized solutions allow proper separation of industrial knowledge, provided that inputs and outputs are clearly defined. This paper builds on the results of [1] and proposes an optimized approach for decentralized fault-focused residual generation. It also introduce the concept of Fault-Driven Minimal Structurally-Overdetermined set (FMSO) ensuring minimal redundancy. The method decreases communication cost involved in decentralization with respect to the algorithm proposed in [1] while still maintaining the same isolation properties as the centralized approach as well as the isolation on request capability.

Spacecraft Attitude Determination System Using Nano-Optical Devices and Linux Software Libraries

This paper presents a new spacecraft attitude determination system based on small optical devices and Linux-based software. This technology intends to support nanosatellite operations by providing low-cost, low-mass, low-volume, low-power, and redundant attitude determination capabilities with quick and straightforward onboard programmability for real-time spacecraft operations. The chosen commercial-off-the-shelf optical devices perform sensing and image processing on the same circuit board, and they are biologically inspired by insects’ vision systems, which measure optical flow and/or track objects while navigating the environment. The firmware on the devices is modified here to enable communication with PC/104 form-factor embedded computers running RealTime Application Interface for Linux. Algorithms are developed for operations using optical flow mode and point-tracking mode, and an application programming interface, along with Simulink® S-functions, is created. The performances of the proposed system, used in optical flow mode, point-tracking mode, and a combination of the two, are assessed using a spacecraft simulator at Rensselaer Polytechnic Institute, and they are compared with measurements from the PhaseSpace® motion-tracking system.

A Correction Algorithm of Aberration of Light for Spacecraft Attitude Determination System

There is the long periodicity attitude error between true attitude and measurement attitude using star sensor for spacecraft attitude determination system because of aberration of light. Aberration of light occurs because the spacecraft’s velocity has a component that is perpendicular to the line traveled by the light incoming from the star. The type of aberration is analyzed and their constants of aberration are calculated in this paper. According to the constants the aberration, the correction mathematical models of parallax of aberration of light of these types of aberration are derived. The parallax of aberration of light of the recognized stars in the FOV of star sensor is calculated with the mathematical models. Then the true vectors of recognized stars at image space coordinate system of star sensor are calculated. The measurement attitude of star sensor is calculated with the true vectors of recognized stars and their vectors at celestial sphere coordinate system. The simulations show the long periodicity attitude error is corrected with the method in this paper. At last the correction of aberration of light was successfully demonstrated using two star sensors with real sky experiment in 2011.

Fault detection and diagnosis in a spacecraft attitude determination system

This paper presents a new scheme for fault detection and diagnosis (FDD) in spacecraft attitude determination (AD) sensors. An integrated attitude determination system, which includes measurements of rate and angular position using rate gyros and vector sensors, is developed. Measurement data from all sensors are fused by a linearized Kalman filter, which is designed based on the system kinematics, to provide attitude estimation and the values of the gyro bias. Using this information the erroneous sensor measurements are corrected, and unbounded sensor measurement errors are avoided. The resulting bias-free data are used in the FDD scheme. The FDD algorithm uses model-based state estimation, combining the information from the rotational dynamics and kinematics of a spacecraft with the sensor measurements to predict the future sensor outputs. Fault isolation is performed through extended Kalman filters (EKFs). The innovation sequences of EKFs are monitored by several statistical tests to detect the presence of a failure and to localize the failures in all AD sensors. The isolation procedure is developed in two phases. In the first phase, two EKFs are designed, which use subsets of measurements to provide state estimates and form residuals, which are used to verify the source of the fault. In the second phase of isolation, testing of multiple hypotheses is performed. The generalized likelihood ratio test is utilized to identify the faulty components. In the scheme developed in this paper a relatively small number of hypotheses is used, which results in faster isolation and highly distinguishable fault signatures. An important feature of the developed FDD scheme is that it can provide attitude estimations even if only one type of sensors is functioning properly.

Observability of a Geosynchronous Spacecraft Attitude Determination System

The observability is examined of a standard geosynchronous spacecraft attitude determination system that uses Earth and sun sensor measurements. The results show that the attitude determination error state, including the attitude and gyro bias errors, is entirely observable over the portion of the orbit where both sun and Earth sensor measurements are available. However, when the sun is outside the sun sensor window, the system is unobservable with a one-dimensional unobservable subspace. It is shown that the eigenvalue associated with this unobservable subspaceiszero and thatthecorresponding eigenvectorisanequilibrium lineoftheattitudedetermination system. The state vector converges to this equilibrium line from any initial state. This result has practical importance in spacecraftattitudecontrolandmaybeusedtospeedtherecoveryfromattitudedeterminationanomalies.Numerical simulations and actual spacecraft e ight data are included to illustrate the theory.

GPS code and carrier multipath mitigation using a multiantenna system

Multipath is a major source of error in high precision Global Positioning System (GPS) static and kinematic differential positioning. Multipath accounts for most of the total error budget in carrier phase measurements in a spacecraft attitude determination system. It is a major concern in reference stations, such as in Local Area Augmentation Systems (LAAS), whereby corrections generated by a reference station, which are based on multipath corrupted measurements, can significantly influence the position accuracy of differential users. Code range, carrier phase, and signal-to-noise (SNR) measurements are all affected by multipath, and the effect is spatially correlated within a small area. In order to estimate and remove code and carrier phase multipath, a system comprising a cluster of five GPS receivers and antennas is used at a reference station location. The spatial correlation of the receiver data, and the known geometry among the antennas, are exploited to estimate multipath for each satellite in each antenna in the system. Generic receiver code and carrier tracking loop discriminator functions are analyzed, and relationships between receiver data, such as code range, carrier phase, and SNR measurements, are formulated and related to various multipath parameters. A Kalman filter is described which uses a combination of the available information from the antennas (receivers) in the multiantenna cluster to estimate various multipath parameters. From the multipath parameters, the code range and carrier phase multipath is estimated and compensated. The technique is first tested on simulated data in a controlled multipath environment. Results are then presented using field data and show a significant reduction in multipath error.

Precision Spacecraft Attitude Estimators Using an Optical Payload Pointing System

The advancement of satellite program technology mandates that issues affecting future spacecraft system development cost be resolved, while continuing to meet smaller size, lower power, and more stringent mission requirements. In the specie c area of the spacecraft attitude determination system, an optical payload pointing system can servea dual purpose: to support the payload mission and to providepreciseattitude information for the spacecraft functions. As a result, star trackers orotherprecisereferencesystems can beeliminated to reduce the development cost. Additional benee ts include performance robustness to spacecraft motion or disturbance and potential cost saving due to weight reduction, e.g., launch cost. Attitude estimators using an optical payload pointing system and strap-down gyros are derived in detail for two distinct formulations: spacecraft body and inertial formulations. Alternative Kalman e lter equation derivations are shown using the skew symmetric matrix properties of the attitude kinematics equation. It is also shown that both formulations produce identical solutions. However, the body formulation is conceptually easier to understand, and the inertial formulation requires less processing time. To show thespacecraft attitudeestimatorperformanceusing an optical payload pointing system, an exampleisshown for an agile low-Earth-orbit satellite.

Clementine attitude determination and control system

The Clementine spacecraft, launched in January 1994, completed a two-month lunar mapping mission, recording high-resolution multispectral images of the lunar surface in unprecedented detail. Mission requirements on the attitude determination and control system included knowledge to within 0.03 deg and control to within 0.05 deg. The typical spacecraft attitude determination system, employing expensive narrow-field-of-view star cameras, highly accurate inertial measurement units, and assorted coarse attitude acquisition sensors, was replaced by an all-new lightweight state-of-the-art sensor suite to meet low-cost, low-weight, and compressed schedule objectives. The sensor suite, which consisted of two wide-field-of-view star cameras and two inertial measurement units incorporating fiber-optic and ring-laser gyroscopes, was delivered in nine months and provided a weight saving of 90% and a cost saving of 75% in comparison with similar systems. A newly developed lightweight reaction wheel, with a weight saving of 50% in comparison with wheels of similar momentum storage capability, was used for fine attitude control. The Clementine mission provided the first space flight for each of these components. Only two anomalies were experienced: a large drift in gyroscope bias along one axis, and a nondestructive single-event latchup in one inertial measurement unit.

Star trackers, star catalogs, and attitude determination - Probabilistic aspects of system design

Optimizing spacecraft attitude determination systems that use onboard star trackers requires analysis and evaluation of some probabilistic aspects of system design. This paper discusses methods of constructing or compiling optimum star catalogs, which are defined as uniform distributions on a sphere. Both local and global measures of uniformity on a sphere are defined. Application of these methods and measures to a specific problem is also discussed. In addition, Poisson models of star tracker acquisition probabilities are formulated to provide a useful analytical basis for designing and optimizing attitude determination systems. These analytical models and methods lead to rapid and realistic quantitative results and should therefore facilitate making system performance trades. The use of such methods should also reduce the need for performing tedious computer simulations to obtain analogous results. TTITUDE determination of both Earth-orbiting satel- lites, as well as planetary spacecraft, can be performed autonomously using one or more onboard star trackers. An essential part of such an attitude determination system is the onboard star catalog, and the performance of the attitude determination system is closely linked to the characteristics of this catalog. In particular, the total number of stars in the catalog and their spatial distribution largely determine the probability of a given number of stars appearing in a tracker's field of view (FOV); this includes the possibility of there being zero stars from the catalog in the FOV. The purpose of this paper is to apply some analytic methods of probability theory to the following problems: generating optimum star catalogs of any given size and, given such a catalog, evaluating some of the characteristics of the star cata- log/star tracker interaction that are of particular interest when designing attitude determination systems. Evaluating such characteristics will entail some analytical modeling. We shall formulate some simple analytical models that lead to excellent estimates of such quantities as star acquisition probabilities as a function of the number of stars in the catalog, number of operating star trackers, and size of the FOV. Another kind of performance parameter that can be analytically estimated is the duration of intervals during which none of the stars in the catalog appears in a star tracker's FOV. These analytical mod- els should be useful for quantitatively evaluating design and performance trades involving these variables, and for star tracker system engineering in general. Their use may obviate doing tedious computer simulations to obtain quantitative esti- mates of such parameters. Although some related efforts have been reported,1 such efforts have focused more strongly on star tracker hardware than analytical aspects of performance.

The measurement of a large antenna using a spacecraft as a receiver

Two computational methods for obtaining the gain pattern of a large antenna using power levels recorded by a spacecraft are developed from the transmission equation. Nominally, the large antenna is linearly polarized while the spacecraft's receiving antennas are circularly polarized. Since the spacecraft has no attitude control system, the orientation of the receiving antennas is not known a priori nor its orientation the same from orbit to orbit. The first method relies on an independent spacecraft attitude determination system to furnish the orientation of the receiving antennas. It combines this information with receiving-antenna gain measurements to give the power level of the incident radiation. Losses due to polarization mismatch are taken into account, but polarization loss factors are not actually computed. The second method relates simultaneous measurements made by separate channels with separate receiving antennas. By solving a system of equations numerically, one can determine the direction of the incident radiation as well as its power level. This alleviates the need for an attitude determination system. These methods originated in the operations of the spacecraft NUSAT1 (northern Utah Satellite). >

Star-Sensor-Based Satellite Attitude/Attitude Rate Estimator

Recent advancements in star sensor technology suggest that implementations of spacecraft on-orbit attitude determination and control systems based solely On star sensor measurements may soon become practical. This paper defines the requirements such applications impose on the star sensors and the algorithms used to estimate attitude and attitude rate. A practical filter implementation is described. Its open-loop performance is evaluated, and simulation results are presented which suggest that performance consistent with a broad class of spacecraft applications is indeed achievable.