analysis model. The process for building the ADAMS model was the crawl, walk, run approach. A singlearray model was generated first, which was correlated to single-array proof of concept tests. Data from  those tests were used to adjust the ADAMS model so the observed test dynamics correlate to the model.  This proved to be a very iterative process while the design matured. The key data that was correlated was  the time from first motion to full deployment and the settling time for the deployed array to dampen out and  stop moving. A complete model with four SAs was generated from the single-array model.  The analysis of the four-array model verified that the dynamics of the deployment in the microgravity  environment should be successful and not detrimental to the dynamics of the CubeSat from either the  nominal case, off-nominal cases, or any of the various initial conditions also analyzed. The various initial  conditions analyzed included a no-rotating and free-floating case, moderate rotations (5, 20, 5 deg/s), high  rotations (19, 80, 19 deg/s in the pitch, roll, and yaw directions respectively), and a hung-up array that  impacted at a later time. These cases were sufficient for this CubeSat project, but for a more critical project,  a full Monte Carlo analysis should be performed to capture more cases that may exist. An interesting result  from the analysis was for the case when the CubeSat was tumbling or rotating; the deployment will  somewhat reduce the rate of rotations while stabilizing the spacecraft.  Flight Assembly and Test  Building and testing the flight hardware started in February 2017. A number of unexpected issues arose  during the assembly, which emphasized the need for schedule margin to work through those issues.  Several of the hardware components required rework upon receipt from manufacturing. The R&R failed  during subcomponent testing, primarily due to incorrectly-sizing the linear actuator and not fully  understanding the friction forces in the pin-puller. This resulted in redesigning the R&R’s 2nd stage  compression springs and reducing their spring rate and stowed compression force. During the hinge  assembly, one of the SMA springs fractured and had to be replaced. Then, during a critical time of final  CubeSat integration with only a month left before the project was scheduled to be completed, multiple SMA  hinge springs fractured and failed prematurely as seen in Figure 6. All of these issues did not occur during  the EDU assembly and test.  Figure 6. SMA Hinge Spring Failed Flight Attachment  The failure of the SMA hinge springs during the final moment of integration proved to be a major issue and  adversely impacted the schedule. From the failure investigation, it was determined that the failure occurred  due to fatigue and over constraining the SMAs at the attachment point on the deployable SA panel. The  EDU was cycled several times and this failure never occurred, making this failure even more unexpected.  Further comparing the EDU and flight hardware, the issue was determined to be the SMA’s attachment  method to the panels. In the EDU version, there was a gradual bending slope in the SMA sheet where it  attached to the deployable panel. In the flight version, a washer under a rivet head was added to prevent  the process of driving the rivet from potentially cracking the SMA sheet, and the SMA was soldered more  firmly past the rivet head toward the hinge pin. It was intended to solder the EDU hinges following the same  procedure used in flight hardware. However, the EDU never adhered to the deployable panel near the rivet,  and testing was conducted in that manner. This was corrected in the failed flight version by having a trained  technician perform the soldering procedure, which was thought to have made the attachment better. The  result actually caused the SMA to bend and crease at a very sharp corner when in the deployed  NASA/CP—2018-219887 386