## 2 Outline of the Rocket and the Launcher
### Engine  
There is a technical issue in hybrid rocket propulsion-- a low fuel regression rate. This can cause low levels of thrust. To overcome this issue, much research has been conducted, and many developments have been made. For example, [PERSON] et al. developed and studied a grain with a star swirl port [12; 13]. [PERSON] et al. developed and studied a grain with a circular helical port [14; 15]. [PERSON] et al. studied the improvement in the regression rate of hybrid rockets using a swirling oxidizer flow [16; 17]. The authors also developed a 3D-printed solid fuel grain with a star-fractal, cherry-blossom-fractal, and star-fractal swirl port [18]. Figure 2 shows that the cross-section of the port has star-fractal geometry, and it is swept while rotating axially. The purposes of the star-fractal swirl port are to have a large surface area and to generate a swirling oxidizer flow. The effectiveness of this port was demonstrated in ground combustion experiments, and the improvement in the regression rate of the fuel was evaluated [19]; hence, this port was adopted for the launch experiment. The fuel grain was made by a fused-deposition-modeling-type, commercial-grade 3D printer. The filament material of the 3D printer was an acrylonitrile butadiene styrene (ABS) resin. Because 3D printing was a time-consuming process, only portions near the \begin{table}
\begin{tabular}{c c c c c} \hline \hline
**Item** & **Unit** & **CAMUI-500P** & \
\begin{tabular}{c} **FY2018 Model,** \\ **“Fukushima-** \\ **Maru”** [8]** \\ \end{tabular} &
\begin{tabular}{c} **FY2021 Model,** \\ **“Amanuma-** \\ **Maru”** \\ \end{tabular} \\ \hline Gross liftoff mass, \(m_{0}\) & kg & 68 & 37.3 & 57.1 \\ \hline Mass of the fuel before & & & & \\ combustion, \(m_{\text{f}0}\) & kg & 7.04 [11] & 3.60 & 7.84 \\ \hline Mass of the fuel after & & & & \\ combustion, \(m_{\text{f}}\) & kg & 1.64 [11] & 0.97 & 1.73 \\ \hline Mass of the oxidizer, & & & & \\ \(m_{\text{ox}}\) & kg & 13.2 [11] & 13.0 & 24.0 \\ \hline Mass of the propellant, & & & & \\ \(m_{p}=(m_{p0}-m_{\text{f}})+m_{\text{ox}}\) & kg & 18.6 [11] & 15.6 & 30.1 \\ \hline Propellant mass & & & & \\ fraction & \% & 27.4 & 41.8 & 52.7 \\ \(\zeta=m_{p}/m_{0}\) & & & & \\ \hline \hline \end{tabular}
\end{table} 
Table 1: Mass properties. Figure 2: Star-fractal swirl-shaped port. port were 3D-printed. Commercially available ABS pipes were used for other portions of the fuel grain. This also led to cost reductions. The detailed manufacturing method was described in reference [19]. The fuel grain was covered by a glass fiber-reinforced plastic (GFRP) ablator and inserted into a motor case. There were two types of motor cases: the external thread type and the welded flange type. These motor cases were different only in the fastening method between the motor case and the injector, and between the motor case and the nozzle housing. For these motor cases, however, their function as a combustion chamber was the same by design. The oxidizer, nitrous oxide (N\({}_{2}\)O), was injected into the combustion chamber through the injector. The bell nozzle was made of graphite and had a 31 mm throat. The development of the engine was started in 2019. The preliminary model was designed, developed, and tested. However, several ruptures occurred in the combustion test. Their cause was determined to be self-decomposition of the oxidizer, nitrous oxide [20], and insufficient strength of the resin housing of the nozzle. Improvements made to both the oxidizer and housing prevented the rupture. Figure 3 shows the combustion test facility with the measuring system, which was also improved to have enough strength to prevent failure on engine rupture. Table 2 shows the parameters of the devices, and measurement uncertainty. Table 3 shows a comparison of total impulses.