Results of the attitude reconstruction for the UniSat-6 microsatellite using in-orbit data

UniSat-6 is a civilian microsatellite that was launched in orbit on the 19th of June, 2014. Its main mission consisted in the in-orbit release of a number of on-board carried Cubesats and in the transmission to the UniSat-6 ground station of telemetry data and images from an on-board mounted camera. The spacecraft is equipped with a passive magnetic attitude control system. Gyros and magnetometers provide the information about the attitude of the spacecraft. The importance of reconstructing the attitude motion of UniSat-6 lies in the dual possibility, for future missions, of:controlling the direction of ejection of the on-board carried satelliteshaving an accurate pointing for remote sensing operation.The reconstruction of the attitude motion of UniSat-6 is based on the data of the on-board Commercial Off The Shelf (COTS) gyros and magnetometers, downloaded at the passages over the ground station in Roma, Italy. At ground, these data have been processed with the UnScented QUaternion Estimator (USQUE) algorithm. This estimator is an adaptation of the Unscented Filter to the problem of spacecraft attitude estimation. The USQUE is based on a dual attitude representation, which involves both quaternions and Generalized Rodrigues Parameters. In this work, the propagation phase of the algorithm contains only a kinematic model of the motion of the spacecraft. This paper presents the results of the reconstruction of the UniSat-6 attitude using on-board measurements. The results show that the spacecraft effectively stabilized its attitude motion thanks to the on-board magnetic devices

Modular Attitude Control System for Microsatellites with Stringent Pointing Requirements

Advancing technology has allowed for the development of low cost attitude control hardware for microsatellites. However, the attitude control design and software development remain a significant cost driver. The Dynacon High Performance Attitude Control system is a modular control system that makes use of reusable algorithm modules enabling the attitude control system to be applied to several different spacecraft missions with very different performance requirements. The High Performance Attitude Control system is described in this paper, and simulation results are shown for a variety of spacecraft

Attitude and In-Orbit Residual Magnetic Moment Estimation of Small Satellites Using Only Magnetometer

Attitude estimation or determination is a fundamental task for satellites to remain effectively operational. This task is furthermore complicated on small satellites by the limited space and computational power available on-board. This, coupled with a usually low budget, restricts small satellites from using high precision sensors for its especially important task of attitude estimation. On top of this, small satellites, on account of their size and weight, are comparatively more sensitive to environmental or orbital disturbances as compared to their larger counterparts. Magnetic disturbance forms the major contributor to orbital disturbances on small satellites in Lower Earth Orbits (LEO). This magnetic disturbance depends on the Residual Magnetic Moment (RMM) of the satellite itself, which for higher accuracy should be determined in real-time. This paper presents a method for in-orbit estimation of the satellite magnetic dipole using a Random Walk Model in order to circumnavigate the inaccuracy arising due to unknown orbital magnetic disturbances. It is also ensured that the dipole as well as attitude estimation of the satellite is done using only a magnetometer as the sensor

Microsatellite Attitude Determination and Control Subsystem Design and Implementation: Software-in-the-Loop Approach

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. In Taiwan, Air Force Institute of Technology (AFIT), dedicating for students to design experimental low earth orbit micro-satellite, called AFITsat. For AFITsat, the operation of the ADCS consists of three modes which are initialization mode, detumbling mode, and normal mode, respectively. 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. This paper will discuss using the SIL validation ADCS function and verify its feasibility

Attitude Determination and Control Subsystem Design for a CubeSat

This project continues the design and testing of the Attitude Determination and Control Subsystem (ADCS) for a nano-satellite. The primary mission objective is solar X-ray spectroscopy using the Sphinx-NG instrument, which requires that the CubeSat fly in a high-altitude, polar, sun- synchronous orbit pointing to the sun with 1-2 degrees of accuracy. The ADCS requires gyroscopes, sun sensors, and a magnetometer for attitude determination. Attitude control is executed using magnetorquers as actuators. This project focused on the analysis of attitude determination algorithms and control policies to select the most efficient and accurate methods. After method selection, simulations of the ADCS were conducted, and research was performed concerning hardware testing for the ADCS

New Approach to Achieving Stand Alone GPS Attitude Determination using Dual Short Baselines for Small-Satellite

This paper proposes a new approach to GPS (Global Positioning System) attitude determination for small satellite application in LEO (low Earth orbit). Prior knowledge of attitude and integer resolution is not required. The methodology of the new approach includes integer ambiguity search, initial estimation of attitude and line bias, attitude initialisation, path difference estimation and fine attitude determination. The observable is the carrier phase difference measurement between two GPS antennas. A dual short baseline (typical baseline length up to 30 cm) is assumed in this research. The key point to initialising attitude is to estimated the attitude of individual baseline vectors with respect to the reference frame. Elimination of integer ambiguity is a simple task. Two set of vectors are required to determine an initial attitude. Once attitude is initialised, an estimation algorithm based on the extended Kalman filter starts to determine the attitude. The integer ambiguities and cycle slips can be resolved properly. The filter now is converged and, fine attitude is estimated. The robustness of the filtering estimator is tested with simulated anomalous conditions

Development Of Attitude Determination For Student Pico-Satellite INNOSAT.

Information of satellite’s attitude is very useful in autonomous satellite control system. For a small satellite particularly, attitude sensor cannot be installed in the satellite because of the limited weight and power consumption

Improved sensor fusion for flying laptop based on a multiplicative EKF

Flying Laptop is a small satellite carrying an optical communications payload. It was launched in 2017. To improve the satellite’s attitude determination, which is used to point the payload, a new sensor fusion algorithm based on a low pass filter and a multiplicative extended Kalman filter (MEKF) was developed. As an operational satellite, improvements are only possible via software updates. The algorithm estimates the satellite's attitude from star tracker and fibre-optical gyroscope (FOG) measurements. It also estimates the gyroscope bias. The global attitude estimate uses a quaternion representation, while the Kalman filter uses Gibbs Parameters to calculate small attitude errors. Past Kalman filter predictions are saved for several time steps so that a delayed star tracker measurement can be used to update the prediction at the time of measurement. The estimate at the current time is then calculated by predicting the system attitude based on the updated past estimate. The prediction step relies on the low-pass-filtered gyroscope measurements corrected by the bias estimate. The new algorithm was developed as part of a master’s thesis at the University of Stuttgart, where Flying Laptop was developed and built. It was simulated in a MATLAB/Simulink environment using the European Space Agency’s GAFE framework. In addition, the new filter was applied to measurement data from the satellite. The results were used to compare the performance with the current filter implementation. The new Kalman filter can deal with delayed, missing, or irregular star tracker measurements. It features a lower computational complexity than the previous standard extended Kalman filter used on Flying Laptop. The mean error of the attitude estimate was reduced by up to 90%. The low pass filter improves the rotation rate estimate between star tracker measurements, especially for biased and noisy gyroscopes. However, this comes at the cost of potentially less accurate attitude estimates. Educational satellites benefit from the new algorithm given their typically limited processing power and cheap commercial-off-the-shelf (COTS) sensors. This paper presents the approach in detail and shows its benefit

Improvements in Attitude Determination and Control of the Small Satellite Flying Laptop

Precise attitude control is a key factor of many payloads with high ground resolutions, small fields of view or narrow beams such as an optical data downlink. The small satellite Flying Laptop (FLP), launched in July 2017, was developed by graduate and undergraduate students at the Institute of Space Systems of the University of Stuttgart with support by the space industry and research institutions. The satellite is three-axis stabilized with reaction wheels as main actuators. FLP is equipped with the OSIRIS optical data downlink which was built by the German Aerospace Center (DLR). As this instrument is body mounted on an optical bench, the attitude determination and control system (ACS) is required to point the whole satellite in the direction of the ground station with a high pointing accuracy of 150 arcseconds. At the time of launch the ACS did not reach this precision. This paper describes how the attitude determination and control were improved to achieve the required performance. The improvements can be divided into two parts. The first part includes the enhancement of on-board sensor processing and attitude control. In the second part, in-orbit data were utilized to increase the accuracy of parameters which are used to control the spacecraft. The first part includes the addition of a Kalman filter, an improved position propagation, and the introduction of adaptive gains to the on-board ACS. The FLP simulation test bed was used to verify the changes. The test bed was also used to find adequate initial values for the Kalman filter and to find inaccuracies in the sensor processing. In the second part, the adaptive gains and the Kalman initial values were validated in-orbit after the upload of the new sensor processing. Moreover, the on-board component orientation settings were corrected for the star trackers, the multi-spectral camera system, and the OSIRIS instrument on FLP. As a result, the satellite fulfills the pointing requirement of less than 150 arcsecond deviation from the target attitude for a sufficient period of time during a pass over the target. Successful links with the optical data downlink were demonstrated with the DLR ground station in Oberpfaffenhofen

Deterministic and Recursive Approach in Attitude Determination for InnoSAT

Attitude determination system (ADS) was indispensable in attitude control of satellite. Especially for InnoSAT due to the limitation of budget, weight, and power, the attitude was determined using onboard position sensors. Previous research has successfully implemented the attitude determination using only Earth's magnetic field sensors for small attitude angle, but the approach produced quite big error for large attitude angle. This paper presents attitude determination for InnoSAT using combination of sun sensors and earth's magnetic field for large attitude angle. The attitude was determined using a deterministic (QUEST) and recursive (EKF) approach. A problem arises when using the sun sensors while the satellite experiencing eclipse. Consequently, the accuracy of both approaches was analyzed at eclipse and no eclipse conditions. The result shows that deterministic approach produced better accuracy at no eclipse but recursive approach produced better accuracy at eclipse. The strategy to apply the both approaches and eclipse conditions also discussed in this paper