# Novel 3U Stand-Alone CubeSat Architecture for Autonomous Near Earth Asteroid Fly-By
## 3 Autonomous Navigation  
For such compact (in mass and volume) and low-cost configurations, autonomous navigation is required during the mission phases in order to reduce ground tracking demand and to avoid large communication subsystems (MR3 in Table 1). The mission phases defined in Section 2 highlight three different navigation phases:  
1. Earth's relative navigation during ID and EE;
2. absolute heliocentric navigation during DSC, DT, and FD;
3. relative navigation during CE and SA. The first phase can be accomplished (at least partially) by means of Global Navigation Satellite System (GNSS) receivers, which with low volume, mass and power consumption can provide high orbit determination accuracy in Low-Earth Orbit (LEO). Hyperion Technologies GNSS200 [15] is an example of low volume (below 0.001 U), low mass (3 g) and low power consumption (below 150 mW peak) receiver able to provide below 8 m orbit determination accuracy. Clearly, orbit determination accuracy would significantly drop as long as the CubeSat is far from the Earth. Then, during the apogee raising escape trajectory, discussed later in the Propulsion subsection, the orbit determination process via GNSS can be improved by merging with optical navigation, which is exploited in the second phase. Moreover, when the CubeSat is approaching the perigee of the SSGTO, the high orbital \begin{table}
\begin{tabular}{c c} \hline
**Acronym** & **Mission Requirement** \\ \hline MR1 & Total mass shall be below 4 kg \\ MR2 & Total volume shall fit 3U \\ MR3 & The spacecraft shall exploit fully-autonomous navigation strategy \\ MR4 & Mission duration shall not exceed 650 days \\ MR5 & The architecture shall exploit COTS components for all subsystems \\ MR6 & The spacecraft shall be equipped with payload in the visible and IR ranges \\ MR7 & The spacecraft shall be equipped with a propulsion system \\ MR8 & The spacecraft shall be equipped with primary and secondary power systems \\ MR9 & The spacecraft shall be equipped with communication system to downlink the scientific dataset \\ \hline \end{tabular}
\end{table} 
Table 1: Mission Requirements definition. \begin{table}
\begin{tabular}{c c c} \hline
**Phase** & **Acronym** & **Duration** \\ \hline Injection and De-Spinning & ID & \(\sim\)1 day \\ Earth’s Escape & EE & \(\sim\) 3-10 days \\ Deep-Space Cruise & DSC & \(\sim\)200-550 days \\ Close Encounter & CE & \(\sim\)1–14 days \\ Scientific Acquisition & SA & \(\sim\)1–2 days \\ Data Transmission & DT & \(\sim\)2–3 days \\ Final Disposal & FD & \(\sim\)5–6 days \\ \hline \end{tabular}
\end{table} 
Table 2: Description of the mission phases. velocity (\(\sim\)10.5 km/s compared to the \(\sim\)7.5 km/s of a LEO orbit) may influence the orbit determination accuracy. The GNSS receiver orbit determination accuracy evaluation along the escape trajectory, eventually in coupling with optical navigation strategies, is left for future works. The navigation accuracy at the end of the escape trajectory influences the initial conditions for the second navigation phase. To compensate the currently not available accuracy, large initial conditions uncertainties in the second phase have been considered. In deep-space missions, the second navigation phase usually relies on ground tracking. The only two CubeSats which flown in deep-space as of today, JPL MarCO-A and -B, carried on-board JPL Iris transponder, and the navigation team exploited both ranging and Doppler tracking for position estimate [16]. This well proven technique however does not fit a compact 3U configuration in terms of mass and volume, and in terms of cost since it would rely on ground tracking. The optimal solution for such configuration would be the use of already on-board instrumentation, adapted for autonomous navigation purposes.