## 2 Materials and Methods--Implementation, Integration, and Testing of the IDEASSat Flight Model
### Environmental and Functional Test
#### 2.5.3 Vibration Test  
Per launch provider requirements [24], IDEASSat was subject to vibration testing [26], which consisted of a sine burst (quasi-static load) test [27] and random vibration test [28] on all 3-axes. The specifications of the sine burst test are shown in Figures 10 and 11, which show the g-force acceleration of the vibration table as a function of time for the spacecraft Z-axis and X/Y-axes, respectively. The acceleration is in units of g = 9.8 m s\({}^{-2}\). In order to reach the 100% vibration level shown in these figures, the amplitudes of the vibration table had to be sequentially ramped up from 25%, 50%, and 75% levels. As a result, the spacecraft was actually subjected to four sine bursts of increasing amplitude along each axis. It can be seen that per launch provider requirements, the Z-axis was subject to a maximum g-force load of 9.369 g, compared to 14.99 g along the two other axes. The specification of the random vibration test is shown in Table 10, which shows the profile of the power spectral density envelope at different frequency bands between 20 and 2000 Hz. These test levels were defined according to launch provider protoqualification requirements, which specified a 3 dB margin above the maximum predicted environment (MPE) [24]. During the random vibration test, the spacecraft was subjected to a spectrum of multiple vibration frequencies defined by this envelope with the overall g-force acceleration level defined by the GRMS (root mean square) being 7.75 g. The duration of the test was 1 min along each of the three axes of the spacecraft. During testing, the spacecraft was enclosed inside a COTS ISISpace 3U TestPOD approximating the mechanical interface between the spacecraft and the deployer, as shown in Figure 12[29]. The sine burst and random vibration tests on each axis were preceded by a resonance survey to identify the normal modes of the spacecraft. These results were compared to a resonance survey performed after the two vibration tests, with the pass criteria being shifts in the normal mode frequencies of less than 10%, which was successfully attained. Figure 10: Z-axis sine burst test specification of g-force acceleration (in units of g = 9.8 m s\({}^{-2}\)) with respect to time. \begin{table}
\begin{tabular}{c c c c} \hline \hline
**Beacon Format** & **Short** & **Lite** & **Normal** \\ \hline Full Packet Loss (\%) & 0 & 0 & 0 \\ Partial Packet Loss (\%) & - & 50 & - \\ Packet Error (\%) & 56.5 & 100 & 100 \\ \hline \hline \end{tabular}
\end{table} 
Table 9: 40 km end-to-end test result. After the first vibration test of the IDEASSat flight model, the spacecraft showed no signs of visible damage, it was found to power up normally, and functional verification was successfully completed. However, when the UHF transceiver was damaged due to poor impedance matching, the spacecraft had to be partially disassembled for de-mating, and was subjected to a second vibration test following re-integration. During the second vibration test, a single nut used to secure a connector inside CIP separated, due to lack of epoxying. After the functional testing, it was found that although the separated nut did not affect the function of CIP, partial disassembly would again be necessary to retrieve the separated component to mitigate the risk of possible short circuit. During this process, the nut was reattached and properly epoxied. A survey was also performed to verify that all fasteners and connectors were properly epoxied or staked. After discussions and evaluations by the launch provider, the spacecraft team, and the payload team, it was determined that a third vibration test would not be necessary as the noncompliant fixture had been corrected and the spacecraft was verified to be functional.