Attitude Determination System Research Articles

Model-Based Design and Testbed for CubeSat Attitude Determination and Control System with Magnetic Actuation

Model-Based Design and Testbed for CubeSat Attitude Determination and Control System with Magnetic Actuation

Remote sensing of electron precipitation mechanisms enabled by ELFIN mission operations and ADCS

The Electron Loss and Fields INvestigation (ELFIN) mission comprising two 3U+ CubeSats was developed, built, and operated by several generations of undergraduate students at UCLA. The spin-stabilized CubeSats (spin-rate: 21 RPM) produced high-resolution measurements of precipitating, trapped, and backscattered fluxes of electrons and ions in the radiation belts. Launched in September 2018, ELFIN operated successfully until its deorbit just over four years later. Initially, however, mission operations was very challenging and only tapped the full mission potential after a thorough redesign of the operations paradigm. This mid-mission adjustment yielded the higher data downlink volume necessary to acquire a comprehensive data set, therefore enabling ensemble studies with sufficient statistical significance. The new operational framework also led to additional improvements across the mission. Most notably, the Attitude Determination and Control System (ADCS) benefited from more reliable collections of magnetometer data for attitude determination, enabling knowledge and control to <1°. This allowed high pitch-angle resolution (especially within the loss cone) and high energy resolution spectrograms, both powerful diagnostics of radiation belt electron and ion precipitation. This paper highlights the scientific advancements made possible by ELFIN’s efficient mission operations and ADCS design, unique for CubeSats, and emphasizes the role of electron precipitation measurements for future studies in magnetospheric, ionospheric, and atmospheric physics.

Design and simulation of sun position sensors for space applications: A comparative study.

Sun position sensors are used in space applications as part of the attitude determination and control system. The aim of this work is to describe the design process of single and dual axis sun position sensors based on a photodiode array detector and a window to limit and direct the light that reaches the detector. The design process involves choosing the sensor architecture, modeling its output, and then evaluating the model through simulation. Six architectures, with different configurations and geometries for both the detector and window, were modeled and compared. To evaluate the model and performance of the sensor, a program was developed, and simulations were made varying the window height and size as well as the photodiode size. From the simulations performed, we concluded that for each design, the key factors that influence the sensitivity and field of view performance are the window height and window or photodiode size, depending on the sensor configuration. From the comparison across architectures, the detector configuration, as well as the window geometry, influences the sensitivity and linearity of the response for similar conditions of operation. The two-quadrant detector configuration offers a better sensitivity than the triangular photodiode configuration. For a better linearity of the output, a square window is preferred. This comparative study is part of the development of new products for space applications by the National Atomic Energy Commission and contributes to the Argentinian National Space Plan.

The Effects of Pointing Error Sources on Energy Delivery from Orbiting Solar Reflectors

Many proposals are being made for cleaner, more sustainable forms of energy production. Terrestrial solar photovoltaic farms (SPFs) could potentially deliver large quantities of energy to the grid, although these are limited to daytime use. The output from these SPFs could be enhanced, particularly around dawn and dusk, by the use of orbiting solar reflectors (OSRs) in near-polar orbit. These would reflect an image of the solar disk, or solar image (SI), onto the SPFs to augment their energy output. Pointing requirements are therefore to ensure that reflected sunlight is delivered to the terrestrial SPF, avoiding the losses incurred by an offset of the SI and the SPF itself. The SI would typically be of order 10 km for a reflector in a 1000 km orbit. Given the potentially large size of the reflectors, this presents a challenge for the attitude determination and control system (ADCS) to ensure that the maximum quantity of energy can be delivered to a SPF, typically requiring large control moment gyro actuators. In addition, there exist numerous sources of error in the ADCS which can cause further degradation in the quantity of energy delivered to the SPF. These errors can manifest in the resolution of the various sensors, flexible structural modes, manufacturing inaccuracies, and misalignments due to vibration during launch. This paper will investigate the effects of pointing error sources (PES) on the reflector ADCS and so on the quantity of energy delivered to the SPF. With the application of a PD controller with feedforward compensation, and the set of noise characteristics defined in this paper, numerical simulations will show the typical losses in energy delivered to SPFs of 0.015% when the model accounts for PES in onboard sensors, actuator uncertainty and flexible structural modes.

Stochastic analysis of drift error of gyroscope in the single-axis attitude determination

Understanding and mitigating the various stochastic errors in gyroscope measurements is crucial for accurate attitude determination. White noise and bias instability can introduce drift errors, reducing the system’s accuracy. Efficient sensor selection and algorithm design necessitate a thorough understanding of these influences and how they propagate throughout the system. This paper addresses the propagation of significant stochastic errors in the attitude determination system, including white noise (angle random walk), bias instability, rate random walk, and correlated noise. The paper discusses the modeling strategy and statistical characteristics of these error contributors. It then evaluates the variance propagation of these errors within the output of a single-axis attitude determination system. The presented analysis calculates and compares the impact of each source of error on the system’s precision over time. Experiments with the MPU6050 and L3G4200D sensors are conducted to validate the analytical results.

POSMind: developing a hierarchical GNSS/SINS post-processing service system for precise position and attitude determination

POSMind: developing a hierarchical GNSS/SINS post-processing service system for precise position and attitude determination

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.

Satellite position and attitude estimation using an infrared earth sensor

Micro/nano-satellites are constrained by payload, size, and cost, making enhancements in design flexibility crucial, particularly for orbit and attitude determination systems. An effective strategy is the innovative utilization of existing sensors. This study investigates a novel application of Earth sensors for autonomous position and attitude determination enabling them as a viable option for autonomous navigation in low-cost micro/nano-satellites through computer vision technology. To address large resolution differences and distortions between sensed infrared images and reference maps, this paper presents a coarse-to-fine matching framework that integrates deep learning features and area-based registrations. First, CNN-based registration and optimization techniques are utilized for coarse registration. A modified VGG16 (M-VGG) and a multiscale pyramid information fusion (MPIF) module are employed to extract and describe deep abstract features, enhancing robustness. Subsequently, edge-based extended phase correlation is employed for fine registration. Finally, the ultimate position and attitude parameters are determined by integrating the results from coarse and fine registration. The effectiveness of the proposed scheme is validated using synthetic images. An analysis of 342 images demonstrates sub-pixel registration accuracy, with a root mean square error (RMSE) of 1.9322 km for position and 0.1185° for attitude.

A framework for developing an attitude determination and control system simulator for Cubesats: Processor-in-loop testing approach

The outbreak of miniaturized satellites is increasingly demanding well-performing attitude determination and control systems (ADCS). Testing an ADCS directly using hardware is extremely costly and risky, whereas testing it using only computer simulation does not yield realistic enough results. Therefore, a detailed framework for developing an ADCS simulator with processor-in-the-loop testing for low-earth orbit Cubesats is established in this paper. In particular, the study focuses on the implementation of the algorithms on a STM32 with ARM Cortex CPU for efficient testing and prototyping. This is done by formulating a quaternion-based satellite dynamic model equiped with a three-axis magnetorquers and a set of pyramidal cluster of reaction wheels for high pointing capabilities. Prominently used sensors are modeled so that the attitude can be estimated using a multiplicative extended Kalman filter. For attitude control, several modes are developed based on different mission stages, and a simplified less burdensome control switching strategy is proposed to maintain stability with minimal computation and energy resources. To test the applicability of the simulator, an ADCS is designed to meet the requirements of the recent low-atmosphere measurement missions. The results are analyzed in accordance with the ESA ECSS standards, and the assessment shows that the specifications can be met with a safe margin, leaving room for unmodeled errors to be taken into account in later stages of the design. Additionally, the proposed control strategy gives various possibilities for missions with different types of objectives, furthermore the framework's flexibility provides easy modification of any part for the user's specific needs.

Innovative Exploration of a Bio-Inspired Sensor Fusion Algorithm: Enhancing Micro Satellite Functionality through Touretsky's Decentralized Neural Networks

Insect-inspired sensor fusion algorithms have presented a promising avenue in the development of robust and efficient systems, owing to the insects' ability to process numerous streams of noisy sensory data. The ring attractor neural network architecture has been identified as a noteworthy model for the optimal integration of diverse insect sensors. Expanding on this, our research presents an innovative bio-inspired ring attractor neural network architecture designed to augment the performance of microsatellite attitude determination systems through the fusion of data from multiple gyroscopic sensors.Extensive simulations using a nonlinear model of the microsatellite, while incorporating specific navigational disturbances, have been conducted to ascertain the viability and effectiveness of this approach. The results obtained have been superior to those of alternative methodologies, thus highlighting the potential of our proposed bio-inspired fusion technique. The findings indicate that this approach could significantly improve the accuracy and robustness of microsatellite systems across a wide range of applications.

Innovative design and development of attitude determination and control systems for CubeSats with reaction wheels

Attitude determination and control systems (ADCS) represent a critical facet of CubeSat missions, orchestrating the precise orientation and stabilization of these small satellites in the space environment. This paper presents a comprehensive design and development of an ADCS tailored for CubeSats, harnessing a reaction wheel system to deliver a cost-effective and dependable solution for small satellite applications. The research begins by elucidating the requisites and specifications for the ADCS and then delves into the design phase, complemented by intricate modelling and simulation employing MATLAB Simulink and the Webots Simulator. The results of this study underscore the exceptional performance of the proposed ADCS configuration, leveraging the reaction wheel model. This system demonstrates an unparalleled capacity to achieve precise and controlled attitude adjustments, well within the defined parameters. Furthermore, this research underscores the pivotal role played by efficient system design, meticulous simulation, and rigorous testing in the triumphant implementation of ADCS, greatly enhancing CubeSat missions and their contributions to the realm of space exploration and technology innovation. This comprehensive approach to the design and testing of an ADCS for CubeSats ensures that these diminutive satellites continue to make significant strides in space missions, paving the way for an exciting future of space research and technology development.

Attitude maneuver planning and robust tracking control for flexible satellite

Abstract In this paper, we investigate the attitude manoeuver planning and tracking control of the flexible satellite equipped with a coilable mast. Due to its flexible beamlike structure, the coilable mast experiences bending and torsional modal vibrations in multi-direction. The complex nonlinear coupling and other external disturbances significantly impact the achievement of high-precision attitude control. To overcome these challenges, a robust attitude tracking controller is proposed for easy implementation by the Attitude Determination and Control System (ADCS). The controller consists of a disturbance compensator, feedforward controller and output feedback controller. The compensator, based on a Nonlinear Disturbance Observer (NDO), effectively compensates for the cluster disturbances caused by vibrations, environmental factors and parameter perturbations. The feedforward controller tracks the desired path in the nominal satellite model. Furthermore, the output feedback controller enables large-angle manoeuver control of the satellite and evaluates the suppression effect of the controlled output on the observation error of cluster disturbances used the ${L_2}$ -gain. Simulation results demonstrate that the proposed controller successfully achieves high-precision attitude tracking control during large-angle manoeuvering.

An Improved Gaussian Sum Extended Kalman Filter With Colored Noise for GNSS/SINS Tightly Coupled Positioning and Attitude Determination Systems

An Improved Gaussian Sum Extended Kalman Filter With Colored Noise for GNSS/SINS Tightly Coupled Positioning and Attitude Determination Systems

Precision space telescope attitude determination and control system design principles

In order to preclude the impact of atmospheric factors on the quality of received data, state of the art astronomy research projects require spacecraft mounted telescopes. However, such designs entail the issue of sustaining the accuracy of telescope boresight attitude relative to the object of observation for the required period of time. This paper is a summary of design principles for an attitude determination and control system of a space telescope mounted rigidly onboard a spacecraft. The paper formulates requirements for attitude accuracy, and proposes a viable equipment configuration as well as a three- stage attitude control algorithm. The first stage of the algorithm ensures primary boresight pointing towards the target object prior to the object being captured in the telescope’s FoV based on star tracker and angular rate sensor data. The second stage of the algorithm ensures boresight positioning relative to the target object direction up to an error value allowing for the required image quality, based on telescope readings. The third stage of the algorithm ensures telescope boresight holding relative to the target object direction with a required accuracy for the duration of exposure. Interferences from operating satellite equipment are reviewed, and the principles of reducing their impact on sighting accuracy are formulated. Low inclination, low eccentricity geostationary orbit is recommended as the best operating environment for the space telescope’s scientific instrumentation, as well as for providing a continuous data interface between spacecraft and ground segment.

Novel extended Kalman filter + H∞$$ {\mathbf{H}}_{\infty } $$ optimal output feedback control configuration for small satellites with high pointing stability requirements

SummaryOne of the most challenging subsystems for the CubeSats is the attitude determination and control system (ADCS) because it involves hardware and software integration, advanced strategies and algorithms to determine and control the attitude that impacts on space maneuvering and well working of a large set of payload (e.g., the optical payload). Although ADCS is largely studied, it still requires further investigations and analysis in order to achieve high‐performance and a reduced efforts. This paper presents an effective and reliable solution for the ADCS of CubeSats involved in Earth observation missions. The solution is based on an innovative framework that leverages the strengths of two distinct methodologies—H‐infinity optimal output feedback control and the extended Kalman filter (EKF)‐to significantly enhance the pointing stability of satellite systems. While H‐infinity optimal output feedback control is traditionally associated with partial state knowledge, we intentionally extend its application to scenarios with complete state information. The study includes a comparative performance analysis involving three algorithm configurations: the classic combination of EKF and model predictive control (MPC), the utilization of H‐infinity optimal output feedback control without EKF, and our proposed integrated approach. Performance evaluations are based on extrinsic indicators, demonstrating the advantages of our method in terms of pointing stability and overall control system performance. This research underscore the significance of combining innovative thinking with well‐established methodologies, unlocking new possibilities for stability enhancement in a range of engineering applications.

Methodology for CubeSat Debris Collision Avoidance Based on Its Active ADCS System

This research assesses the feasibility of a collision avoidance methodology for CubeSats lacking propulsion. The approach involves altering the satellite’s orientation to modify its cross-sectional area and, subsequently, the drag force. Examining altitudes within low Earth orbit (LEO) across 2U, 3U, and 6U CubeSat formats, maneuvers are considered two days before the Time to Closest Approach (TCA). Evaluation against the Conjunction Data Messages (CDMs) threshold miss distances reveals a minimum 7% and maximum 106% deviation in Vertical Distance Difference (VDD), and 68% to 1045% in Horizontal Distance Difference (HDD) concerning the notification threshold. These findings strongly endorse the practicality of the proposed collision avoidance methodology, utilizing CubeSat Attitude Determination and Control Systems (ADCS). Ongoing research focuses on assessing ADCS maneuver execution rates and implementation times, advancing our understanding and applicability of this innovative CubeSat collision avoidance approach.

Assessment of magnetic field on body axis and orbit axis of RazakSAT in Near Equatorial Orbit using IGRF model

Magnetometer is reference sensor of satellite that function to measure the vector of magnetic field during satellite enter the orbit. The magnetic field measurement vector will be employed in the Attitude Determination System (ADS) to calculate the satellite’s orientation with respect to the Earth. However, any misalignments or disturbances on the ADS can affect the orientation of the satellite in terms of the orbital position. Therefore, in this paper assessment of magnetic field on body axis and orbit axis of RazakSAT by using The International Geomagnetic Reference Field (IGRF) model. RazakSAT as references to verify the IGRF model by using residual analysis that percentage error less than 5% requirement from Astronautic Technology Sdn Bhd (ATSB). The result show, Y-axis and Z-axis from orbit frame meet the requirement of ATSB is less than 5%.

Precise orbit determination of Spire nano satellites

Spire Global, Inc. operates a growing fleet of currently more than 100 CubeSats in different low Earth orbits for commercial Earth observation. These satellites are equipped with dual-frequency GPS receivers and an attitude determination and control system, allowing for precise orbit determination. For three different satellites and a time span of six months we analyze the performance and quality of the on-board collected GPS and attitude data and employ it for precise orbit determination using the Bernese GNSS Software and Napeos, two independent state-of-the-art GNSS processing software packages. We describe technical details crucial for POD and present and compare the in-flight calibrated phase center variation maps. Reduced-dynamic and kinematic orbits are then inter-compared between the two software packages as well as to the orbit solutions produced by Spire Global. We report pseudo-range and carrier phase residuals at the level of 3–4 m and 8–9 mm RMS and a good agreement between the reduced-dynamic and kinematic orbits of around 5 cm 3D RMS. The reduced-dynamic orbit positions and velocities produced with the two employed software packages agree in average on the level of 6–7 cm and 0.05–0.07 mm/s 3D RMS, while the comparison to the orbits produced by Spire Global is markedly worse with 27–30 cm and 0.32–0.36 mm/s 3D RMS. The presented results are an encouraging first step towards using GPS data from the Spire constellation for geodetic, geophysical and ionospheric applications.

ROEKF-MPC Estimator for Satellite Attitude and Gyroscope Bias Estimation

Micro-electro-mechanical system (MEMS) based gyroscopes are commonly used for satellite's attitude determination and control system in recent years. They are packaged in a small form factor and have lower power consumption. They provide a low-cost solution to the emerging NewSpace industry. MEMS gyroscopes exhibit time-zero null bias with variation over temperature. To overcome this problem, this paper proposes a technique for self-calibration of gyroscope bias based on reduced order Extended Kalman Filter (EKF) working alongside an estimator based on the model predictive control (MPC) approach. The proposed technique is referred as the ROEKF-MPC Estimator. Both simulation results and experimental verification using an in-house developed spacecraft simulator are presented. Unlike the EKF method, the proposed method can estimate the gyroscope bias instantly and it is robust against changes in temperature. In addition, the proposed method employing a reduced order EKF is 28.4% computationally more efficient than the typical higher order EKF method. Results show that the pointing performance is better than 0.35deg making the proposed method very attractive for most NewSpace applications. While the proposed method allows for self-calibration of gyroscope bias, its performance is affected by the gyroscope noise and requires the satellite's attitude to be nearly in steady state. This limitation will be addressed in detail.

Attitude Determination System with Rotation of Gyroscopes

A method of auto-compensation for the influence of constant biases in the output signals of gyroscopes in the system for determining the orientation of an object, based on the rotation of the gyroscopes relative to the object, is considered. The analysis of the obtained models of errors in determining the orientation and the simulation performed show the high efficiency of this method for improving the accuracy of readings. The proposed error model makes it quite easy to evaluate the efficiency of auto-compensation. It is shown that the problem may be reduced to integrating the direction cosine matrix that determine the position of the sensitive elements in the reference coordinate system.