# Novel 3U Stand-Alone CubeSat Architecture for Autonomous Near Earth Asteroid Fly-By
## 4 Subsystems Analysis
### Propulsion System
#### 4.2.1 Chemical Micro-Propulsion  
NASA-JPL Small Body Database Browser [37] has been used to preliminary compute the required Earth's SOI escape velocity, \(V_{\infty}\), to inject the spacecraft in an interplanetary transfer towards the target NEA. The Mission Design Tool solves the Lambert problem and provides Pork-Chops plots. It allows the user to enter the name or code of a particular asteroid, giving back information on a mission to the specified target, such as required escape velocity, the ToF, and departure date. Table 3 gives a short list of targets with their departure dates, ToF, and required escape velocity, which have been chosen as the minimum \(\Delta V\) solutions, but not all meet MR4. Recalling the SSGTO injection orbit, and using the patched conics approach, it is possible to compute the approximated \(\Delta V\) to deliver to the spacecraft at the perigee, to inject it on an escape hyperbola:  
\[\Delta V=\sqrt{V_{\infty}^{2}+\frac{2\mu}{r_{SSGTOp}}-V_{SSGTOp}}, \tag{3}\]  
where \(V_{\infty}\) is the hyperbolic excess velocity, \(\mu\) is the Earth gravitational parameter, and \(r_{SSGTOp}\) and \(V_{SSGTOp}\) are respectively the radius and velocity at the perigee of the SSGTO. Based on Equation (3), Figure 3 shows the relation between the escape velocity \(V_{\infty}\) and the perigee \(\Delta V\). As it is shown in the plot, 385 m/s of \(\Delta V\) budget will ensure a hyperbolic excess velocity higher than 0.6 km/s. Considering that 0.6 km/s is the largest escape velocity considered in Table 3, 385 m/s can be assumed as a general \(\Delta V\) requirement for the escape trajectory, the TCM(s), and wheels de-saturation. TCM has been described in the Navigation section, while wheels de-saturation will be discussed in the ADCS section. Further studies can be done in order to define the optimal escape strategy, which due to the limited burning time of the micro-propulsion system, can be fractionated in sequential apogee raising maneuvers. \begin{table}
\begin{tabular}{c c c c c} \hline \hline
**Target** & **Depature Date** & **ToF [days]** & \(V_{\infty}\) [km/s] & **Diameter [m]** \\ \hline
99942 Apophis & 6 August 2028 & 230 & 0.1 & 320 \\
4660 Nereus & 18 May 2028 & 100 & 0.2 & 330 \\
35107 & 14 January 2025 & 205 & 0.2 & 929 \\
161989 Cacus & 17 December 2039 & 635 & 0.2 & 635 \\
153201 & 1 July 2032 & 535 & 0.3 & 510 \\
66391 Moshup & 1 February 2036 & 115 & 0.3 & 1317 \\
13651 & 3 April 2029 & 815 & 0.4 & 562 \\
7482 & 8 April 2031 & 640 & 0.4 & 1052 \\
4034 Vishnu & 27 April 2033 & 510 & 0.4 & 420 \\
142464 & 20 August 2034 & 625 & 0.5 & 886 \\
138127 & 24 May 2027 & 260 & 0.5 & 754 \\
65679 & 17 November 2036 & 615 & 0.5 & 918 \\
1943 Anteros & 20 December 2025 & 165 & 0.5 & 2300 \\
2102 Tantalus & 27 April 2034 & 635 & 0.5 & 1650 \\
3200 Phaethon & 4 March 2028 & 985 & 0.6 & 300 \\
11500 Tomaiyowit & 12 April 2035 & 235 & 0.6 & 738 \\ \hline \hline \end{tabular}
\end{table} 
Table 3: Target NEAs characteristics. Having defined a general \(\Delta V\) requirement, the micro-propulsion system can be evaluated. Solid propellant thrusters are not considered, because they would allow one (or few) burn only, if no customized solutions are designed. On the other hand, cold gas micro propulsion systems are characterized by low specific impulse. Therefore, liquid mono- and bi-propellant micro-propulsion systems represent the best solution to inject the CubeSat to a transfer trajectory towards a NEA. On the market, both single thrusters and complete integrated micro-propulsion systems are available. The former allows a more customized sizing in terms of propellant mass, while the latter allows the integration of a ready-to-use micro-propulsion system.