Redundant push-off springs mounted on COTS bushings are configured inside the cone to give the panel  an initial separation. A COTS microswitch (~$2) is housed behind one of these push-off plungers and used to confirm separation of the hold down. A similar microswitch used on the hinges then confirms full deployment and latching. The solar arrays use triple-junction gallium arsenide solar cells which are in  close proximity of the HDRM in order to achieve maximum panel coverage. The induced shock from the HDRM bolt fracture was a c oncern on the micrometer-thick fragile cover glass on the cells, in addition to  the hinge’s deep groove radial ball bearings and other electronic components within the spacecraft. To  reduce the shock, thin film aluminum honeycomb core material was located at each end of the Frangibolt which serves to absorb the shock forces when the bolt fractures.  A ¼-inch Frangibolt was selected for the HDRM and is assembled into a removable cartridge which  attaches to the cup. The Frangibolt passes through the cup (craft side) and protrudes through the mating  cone (array side) to allow the array to be attached. The cup-cone is thus fastened together via the Frangibolt, and a preload is applied to the front side via torqueing a high-tensile nut up to ~5 N-m. The preload of this bolt is critical to prevent the cup-cone slipping and gapping. In order to prevent cold welding during launch, the Titanium 6Al-4V cup is coated with Titanium Nitride,  while the Titanium 6Al-4V cone is left un-plated. To help accommodate the alignment and assembly constraints at spacecraft level, a design trade was made and it was decided that the cup-cone interface  did not bear on the cone angle itself as in a typical cup-cone interfaces. Instead, the cup-cone elements  bear on flat portions of the corresponding elements with a very small gap (<0.2 mm) to the cone interface  so if slipping did occur, it would then engage. The practical limit of tensile load capability of the Frangibolt is ~6.5 kN. If a larger preload is applied on the cone-side over the maximum in the actuator section, there is risk of creating a gap which could risk  non-fracture of the bolt. There was no space in the design to implement a load cell to give accurate  preloads so the torque scatter was accommodated in the design. Early testing did lead to further  calibration of the test up and an increase in this value to 6.5 N-m as slipping did occur on the engineering qualification model (EQM) vibration test which was evident by visible fretting (Figure 7), but it should be  noted that this did not prevent any deployments.  Figure 7. Engineering Qualification Model cup / cone interfaces showing fretting Early on in the design process of the hold downs it was acknowledged that the temperature of the device  is critical to its operation. Some work was undertaken to better understand the upper temperature limits of  the SMA in the Frangibolt as thermal mission predictions show high temperatures in excess of 80-90°C  during early phases of the missions prior to array deployment. This environmental temperature could result in premature array deployment. It was during this testing that one bolt failed to fracture, upon disassembly there was nothing obvious to show why and the actuator in question subsequently successfully fractured a bolt under normal actuation. The full root cause was never clearly identified but  426