# An Effective Sensor Architecture for Full-Attitude Determination in the HERMES Nano-Satellites
## 2 Effective Sensor Architecture  
The effective sensor architecture for full-attitude determination in the HERMES nano-satellites was developed starting from the imposed mission requirements and system constraints. In addition to those described in the previous section, the system is required to determine its full orbital and attitude states in any point of the orbit, and it shall be capable of withstanding failures in the hardware components. For these reasons, the sensor architecture was selected to be operative both in sunlight and in eclipse and with a few degrees of hot redundancy. First, the mass, power, and cost limitations indicated discarding the star sensors as the primary attitude determination sensors. The possibility to host this sensor typology in a redundant configuration is extremely remote for a nano-satellite. Moreover, the attitude determination accuracy requirement is not extremely strict to force the design to use star sensors. Then, since the attitude state estimation shall not be dependent on the relative attitude with respect to the Earth, the primary attitude sensors were selected to be Sun sensors and magnetometers. The latter are also needed to actuate magnetic torquers, which are very common in nano-satellites. The maximum attitude estimation error drove the selection of fine Sun sensors with maximum measurement errors in the order of \(\sim\)1 deg. However, to avoid the presence of too many fine sensors for redundancy purposes, the sensor architecture was based on a combination of fine Sun sensors in the Sun's primary axes and of distributed photodiodes to retrieve a coarse Sun direction. Note that the photodiodes shall also guarantee the detection of the Sun with a full-sphere field of view, with no blind spots. The requirement to determine the full-attitude state in any section of the orbit imposed the inclusion of gyroscopes for dead-reckoning state propagation while in eclipse. The gyroscope measurements shall be continuously calibrated on-board in order to minimize attitude drifts due to the gyroscope's bias. Analogously, to keep the desired performance, the magnetometers and the Sun sensors are also calibrated with on-board techniques. Specifically, the magnetic sensor's measurements shall be corrected for bias and scale factor errors, while the fine Sun sensors' measurements are adjusted for the presence of Earth's albedo. This sensor architecture requires also the estimation of the orbital position state. This is needed to feed the reference environment models for full-attitude determination and for sensors' calibration. Moreover, this is needed in guidance functions, which need to compute reference attitude states relative to the Earth. Finally, it is fundamental to achieve the mission objectives with the imposed accuracy requirement also on the orbit determination. In the presented architecture, the primary sensor is a GNSS receiver. For redundancy purposes, the GNSS measurements are combined in an EKF with an on-board SGP4 orbital propagator [24] and with the accelerometers' measurements. The former can be updated with a two line element (TLE) set received from the ground. The latter were included because the accelerometers are typically embedded in MEMS inertial measurement units (IMUs) hosting gyroscopes and magnetometers, and they also enable the possibility to deal with orbital control maneuvers. In any case, the measured non-gravitational perturbation accelerations, despite being very disturbed due to the noise level of typical tactical-grade MEMS accelerometers, allow slightly improving the GNSS-based orbit determination accuracy, and they allow increasing the reliability of the on-board orbital propagation [7]. The overall sensor architecture is reported in Figure 1 together with the attitude determination blocks. In the scheme, the interfaces between the different elements and the hot redundancies are graphically reported.