## 5 Results
### Feasibility Analysis of Mission Decision Making  
The training results of the mission decision-making model using RBFNN are shown in Figure 6. \begin{table}
\begin{tabular}{c c} \hline
**Parameters** & **Value** \\ \hline Configuration & Two stages + boosters (liquid) \\ \hline Takeoff mass (t) & 600 \\ \hline Average thrust of the boosters (kN) & 120 \(\times\) 4 \\ \hline Average thrust of the first stage (kN) & 70 \(\times\) 4 \\ \hline Average thrust of the second stage (kN) & 18 \(\times\) 4 \\ \hline Fuel specific impulse of the boosters (m/s) & 2942 \\ \hline Fuel specific impulse of the first stage (m/s) & 2942 \\ \hline Fuel specific impulse of the second stage (m/s) & 3342 \\ \hline Aerodynamic reference area (m\({}^{2}\)) & 10 \\ \hline \end{tabular}
\end{table} 
Table 1: General parameters of the launch vehicle. \begin{table}
\begin{tabular}{c c} \hline
**Parameters** & **Value** \\ \hline \multirow{3}{*}{SSO} & Orbital height (km) & 500 \\ \cline{2-3}  & Orbital inclination (\({}^{\circ}\)) & 97.4 \\ \cline{2-3}  & Orbit-injection mode & Direct orbiting \\ \hline \multirow{3}{*}{Low-inclination LEO} & Orbital height (km) & 500 \\ \cline{2-3}  & Orbital inclination (\({}^{\circ}\)) & 40 \\ \cline{2-3}  & Orbit-injection mode & Direct orbiting \\ \hline \multirow{3}{*}{Elliptical orbit} & Orbital height (km) & 400 \(\times\) 600 \\ \cline{2-3}  & Orbital inclination (\({}^{\circ}\)) & 97.4 \\ \cline{1-1} \cline{2-3}  & Orbit-injection mode & Direct orbiting \\ \hline \multicolumn{3}{c}{Launch site} & Wenchang, China \\ \hline \end{tabular}
\end{table} 
Table 2: Standard mission. \begin{table}
\begin{tabular}{c c c c c} \hline
**No.** & **Failure** & **Failure** & **Orbit-Injection** & **Payload** \\  & **Moment/s** & **Magnitude/\%** & **Target Orbit Type** & **Mode** & **Compositions** \\ \hline I & 200 & 25 & Low-inclination LEO & Direct orbiting & Integral \\ II & 300 & 50 & SSO & Transfer orbiting & Detachable \\ III & 400 & 50 & Elliptical orbit & Direct orbiting & Integral \\ IV & 500 & 75 & Elliptical orbit & Transfer orbiting & Detachable \\ \hline \end{tabular}
\end{table} 
Table 3: List of fault conditions. For the given four groups of fault cases, the mission decision making and trajectory planning are performed using the trained RBFNN and the ATU-IGM, and the simulation results are obtained as shown in Figures 7-11. Figure 6: RBFNN training residual curve. Figure 7: Engine thrust. Figure 8: Flight trajectory and terminal orbit. Figure 9: Perigee altitude. Figure 11: Mass. Figure 10: Apogee altitude. For the transfer orbit entry form, this paper calculates the fuel consumption mass by the TCFC and then leaves the 3% fuel margin for the actual orbit transfer and takes the perigee height and apogee height of the final target orbit as the result of the mission planning and the entry deviation into the transfer orbit as the accuracy of trajectory planning. The results of mission decision making and trajectory planning can be obtained as shown in Table 4.  
The above results show that the proposed method can plan a reasonable and feasible new mission under different fault conditions and mission attributes. For fault I, the launch vehicle lost 90 km at perigee height, but all payloads (9.00 t) entered the mission orbit with sufficient accuracy. For fault II, the launch vehicle firstly entered 410 km \(\times\) 490 km transfer orbit and finally entered the 490 km SSO due to the large magnitude and late occurrence of the fault; the decision-making model enabled the launch vehicle to complete high-precision orbit-injection when it was as close to the original target orbit height as possible and retained most of the payload. For fault III, the launch vehicle lost 50% thrust at the later fault time; the apogee height of the mission orbit decreased by a small amount due to the height limit at fault time and the constraint of the integral direct orbiting, and the carrying capacity could only be matched by significantly reducing the perigee height.