# Novel 3U Stand-Alone CubeSat Architecture for Autonomous Near Earth Asteroid Fly-By
## 4 Subsystems Analysis
### Propulsion System
#### 4.2.2 Electric Micro-Propulsion
Then, now that the approach is 'preliminary' optimized, it is possible to perform a grid search based on 3 varying parameters:  
1. ToF: from 100 to 600 days, with a time step of 10 days;
2. Final velocity: from \(-\)2 km/s to +2 km/s of the fictitious asteroid velocity at the node, with a velocity step of 0.1 km/s (in order to have a wider range of relative velocities);
3. Node passage: the first 10 passages per node (20 in total) after the last ephemerides data available on JPL Small-Body Database Browser (31 May 2020), which translates in around a decade. Then the minimum \(\Delta V\) is found and from the control acceleration profile it is possible to compute the mass and thrust profiles, as a function of the specific impulse (which is assumed constant) and of the BOL mass set as 4 kg, by imposing a set of equations: 2 nd Newton law and mass flow rate equation. Again, the results of this search grid is not necessarily the optimal solution, and it could be improved by extending the final velocity range and by making varying also the in-plane angle. However at this stage it is sufficient to get a rough requirement on propellant mass, that can be eventually increased with a certain margin. The evaluation process has been repeated for the 16 NEAs of Table 3. Table 4 reports the values for ten of them, as the ones with a reasonable propellant mass for a 4 kg CubeSat. The other six (3200 Phaethon, 66391 Moshup, 153201, 142464, 138127, 35107) resulted in extremely large required propellant mass (more than 2.5 kg for a 4 kg CubeSat) and they have been excluded from the table. The other ten showed a maximum required propellant mass of 1.1 kg (for the lowest \(\mathrm{Isp}=2000\) s) and thrust levels within 1.1 mN (with the exception of 4034 Vishnu). Figure 5 reports the extended results for one of the NEAs, 65679. \begin{table}
\begin{tabular}{c c c c c} \hline \hline \multirow{2}{*}{**Target**} & **Mp [g],** & **Mp [g],** & \multirow{2}{*}{**ToF [Days]**} & **Max. Thrust** \\  & **Isp = 2000 s** & & & **[mN]** \\ \hline
99942 Apophis & 650 & 210 & 560 & 0.62 \\
4660 Nereus & 780 & 210 & 510 & 0.98 \\
161989 Cacus & 1100 & 400 & 590 & 1.10 \\
13651 & 900 & 270 & 590 & 0.70 \\
7482 & 900 & 310 & 590 & 1.05 \\
4034 Vishnu & 1400 & 450 & 530 & 1.60 \\
65679 & 830 & 300 & 490 & 0.72 \\
1943 Anteros & 520 & 190 & 490 & 0.52 \\
2102 Tantalus & 850 & 300 & 480 & 1.00 \\
11500 Tomaiyowit & 1100 & 400 & 580 & 1.05 \\ \hline \hline \end{tabular}
\end{table} 
Table 4: Low-thrust analysis results. Considering electric propulsion systems with specific impulse below 2000 s would impact significantly on the propellant mass, affecting the feasibility of the mission. So, Sitael HT-100 [50], Mars Space LTD PPTCUP [51], and Busek BHT-200 [52] are not applicable to this mission. Higher specific impulse options are Busek BIT-3 ion thruster [53], Mars Space LTD mini RF ion engine [54], and Enpulsion IFM Micro 100 Thruster [55] and Nano Thruster [56]. BIT-3 and IFM Micro 100 micro-propulsion systems have however an off-the-shelf configuration too large for a 3U CubeSat, respectively \(\sim\)1.8U and \(\sim\)2.2U, which does not leave enough space for the rest of the subsystems. On the other hand Mars Space LTD mini RF is a thruster occupying approximately 0.6U, but no information are at the moment available for tanks, feeding systems and other components. Required propellant mass for \(\mathrm{Isp}=2500\) s oscillate between 0.5 and 1.2 kg, and considering all the components (tank, feeding system, etc. ), this thruster would require no less than 1.3U. The large volume, in coupling with the required additional design of the rest of the micro propulsion system (as for the ECAPS HGPG highlighted in the previous subsection), makes this thruster not suitable for this application. Enpulsion IFM Nano thruster occupies a relatively small volume (\(\sim\)0.8U), with dry/wet mass 0.67/0.9 kg, ensuring a sufficient propellant mass for almost all the ten reachable NEAs (due to a specific impulse up to 6000 s).