## 6 Space-Segment Optimization Aspects
### Constellation Deployment  
Deployment for an optimal constellation for continuous global coverage requires multiple satellites separated within an orbital plane and distributed over several orbital planes at desired altitudes and inclinations. Deploying a global-coverage constellation with the existing launch paradigms would require a dedicated launch for each orbital plane. This would result in very high launch costs beyond the budget of usual missions [120; 125]. As cost is one of the prime optimization parameters, developing an efficient way to deploy a constellation is essential [126]. This requires addressing the constellation configuration, satellite design, and launch opportunities simultaneously [6]. As presented in [127], constellation deployment  
\begin{table}
\begin{tabular}{l l} \hline \hline
**Launch Vehicle** LV & **Payload Capacity to LEO [kg] \\ \hline Micro-LV & \(\leq\)500 \\ Small-LV & 501-2000 \\ Medium-LV & 2001–20,000 \\ Heavy-LV & \(>\)20,000 \\ \hline \hline \end{tabular}
\end{table}
Table 8: Launch Vehicles Classification.  
may be distributed in direct injection, where Sat is placed by LV, and indirect injection, where either the LV or Sat performs non-planner maneuvers to achieve the final orbit. However, Impulse per unit of Mass (needed to perform a maneuver) (\(\Delta V\)) for such orbital transfers is very high. An alternative indirect injection approach is to utilize orbital perturbations due to Earth's oblateness to transfer Sat different orbital planes and populate the constellation. This idea was first patented by [PERSON] and [PERSON] to use natural perturbations to separate the orbital planes with RAAN. The nodal precession due to Earth's oblateness varies with altitude, inclination, and eccentricity at different rates. The nodal precession using Second-Degree Zonal Harmonic of the Earth's Gravity Field (\(J_{2}\)) is given by Equation (8) [128], where \(\hat{\Omega}_{J_{2}}\) is the rate of change in \(\Omega\) due to \(J_{2}\), \(R_{E}\) is radius of Earth, \(a\) is the semi-major axis, \(e\) is eccentricity, \(i\) is the orbital inclination, and \(n\) is the mean motion of the Sat.  
\[\hat{\Omega}_{J_{2}}=-\frac{3}{2}\frac{R_{E}^{2}}{\left(a(1-e^{2})\right)^{2}} J_{2}n\cos(i) \tag{8}\]  
The plane separation would first require an in-plane maneuver, which requires less \(\Delta V\) to change the orbit for a different nodal precession rate and after the drift period required for desired plane separation maneuver back to the mission orbit [121]. For constellation deployment, [129] uses drag to achieve in-plane maneuvering and nodal precession to achieve out-of-plane maneuvering. FORMOSAT-3/COSMIC satellite mission is presented in [130], where six deployed satellites used in-plane thrust maneuvers for orbit raising and utilized differential nodal precession to achieve different orbital planes. A detailed study to deploy a multi-plane constellation using nodal precession is also presented in [131].  
Using orbital perturbations for plane separation results in a longer time for full constellation deployment, whereas a dedicated launch induces high mission costs, resulting in a trade-off between deployment time and launch cost [125].