# An Effective Sensor Architecture for Full-Attitude Determination in the HERMES Nano-Satellites
## 1 Introduction
It is evident that the usage of extremely accurate sensors is typically not possible, and thus, an effective sensor architecture for full-attitude determination is needed. This paper discusses the one that was developed and verified for the HERMES satellites, but it is possible to state that the presented results are valid for CubeSat applications with the system constraints listed in Table 1.  
Existing literature studies have addressed the problem of sensor architectures for spacecraft attitude determination in real missions for many years [8; 9; 10]. However, only in recent studies have the results been available for nano-satellites' and CubeSats' attitude determination subsystems. The work of [PERSON] presented an attitude determination \begin{table}
\begin{tabular}{l l l l} \hline \hline
**Constraint** & **Value** & **Units** & **Notes** \\ \hline Spacecraft mass & \(\leq\)10 & kg & \\ Spacecraft volume & \(\leq\)6 & U & CubeSat Standard \\ Power budget & \(\leq\)50 & W & \\ On-board computer frequency & \(\leq\)128 & MHz & \\ Memory footprint & \(\leq\)32 & MB & \\ ADCS cost budget & \(\leq\)75 & k\(\epsilon\) & \\ \hline \hline \end{tabular}
\end{table} 
Table 1: Assumed system constraints for the proposed sensor architecture. system based on Sun sensors, magnetometers, and gyroscopes to be processed by an extended Kalman filter (EKF) [11]. The simulations results proved errors in the order of 10 deg to 20 deg, which are acceptable for an experimental pico-satellite in-orbit testing mission, but they are not for missions with imposed scientific requirements. Similarly, [PERSON] presented real flight data from a pico-satellite mission that is capable of estimating the orbital position with an \(\sim\)\(1\times 10^{2}\) km accuracy [12]. A few years later, [PERSON] discussed an attitude determination system that utilizes rate gyros, magnetometers, coarse Sun sensors, and an EKF [13]. In this case, the attitude determination accuracy was about 2 deg to 3 deg in sunlight and decreased to about 7 deg to 8 deg when the states were estimated in eclipse. In the very last years, literature works about the ADCS in nano-satellites became more frequent, and they were often focused on improving the performance and the reliability of the developed system and sensor architectures. [PERSON] proposed a two-stage approach method for attitude determination that was robust to sensor faults [14]. In nominal conditions, this method is accurate in the order of \(10^{1}\) arcsec, but it requires a star sensor, which may not respect the system constraints listed in Table 1. With a different approach, the work of [PERSON] exploited independent vector and gyroscope measurements to efficiently determine the spacecraft attitude [15]. This allowed reaching a high estimation accuracy, but it required a computational load and minimum sensor performances that are not suitable for nano-satellites. In these regards, the problem of attitude determination in nano-satellite was recently investigated by [PERSON] [16], [PERSON] [17], and [PERSON] [18]. These research works exploited miniaturized sensors; however, the obtained accuracy was in the range of 1 deg to 10\({}^{1}\) deg, and they did not provide MIL or HIL test results. Experimental data were exploited in the work of [PERSON] [19], who proposed a simple, yet reliable method to estimate the Sun-pointing direction in a micro-satellite mission. Exploiting solar panels and photodiodes, this work achieved an estimation accuracy in the order of 10 deg. Similar conclusions are also possible analyzing the mission results of the NetSat satellites, whose overview is summarized in the work of [PERSON] [20]. The effective sensor architecture proposed in this paper allows achieving a 1 deg attitude determination accuracy at 1\(\sigma_{a}\) in any section of the orbit, together with a 10 m orbital position estimation at 1\(\sigma_{o}\).