energy needed to be carefully balanced so as not to over-design the parts or cause the brackets to break at the end of deployment latch up. The hinge design is kept generic and latches in the same position, and hence different deployment angles between missions are achieved by simply changing the bracket geometry and not touching the hinge mechanism itself. Each hinge has a simple latch which has a PTFE-coated running surface with a carefully profiled end to give positive locking at the end of deployment.Early versions of these did disengage if enough moment was applied on the hinge so a change was made to bead blast the end surface and after this they never disengaged. The first design used plain bushings; however, the torque of these was too high and varied considerably  over temperature, so this simple and small volume solution was replaced with deep groove radial ball bearings with dry-lubricated PGM-HT cages similar to the ones used in SSTL 10SP reaction wheels.These had a large radial load capacity and very low and consistent running torque (see Figure 5). The initial hinge design used a regular coil torsional spring, but this proved to give unreliable torque performance, and to keep the stresses low in the spring it ended up being quite long and thus not fitting in the design volume on the spacecraft. Instead, a custom beryllium copper clock spring was designed,which had good torque performance through the rotation angle. DMC-CFESAT hinges insitu with platform Details of fixed hingewith microswitch Figure 5. An overview of heritage deploy able array hinges and interface brackets The deployable panel went through an extensive unit-level qualification; the same unit also went through  spacecraft-level testing doubling some of the typical tests performed. Over 20 deployments were performed at ambient pressure cold / hot and well as a Thermal Vacuum Test (TVT) campaign to remove the air drag on the system. During TVT, one test was performed hot with the maximum system energy  and the other cold with one spring removed, the other at low torque and no push-off springs to represent  different possible failure modes and a worst case deployment case; all tests were successful. Original HDRM Development for DMC-CFESAT SSTL started investigating the use of Frangibolts ( see www.tiniaerospace.com) in early 2000’s for use on  its static and multiple panel deployed tracking solar arrays due to the simple design, ease of use, volume, mass and cost of these devices. The development for DMC-CFESAT was particularly mass and volume  critical which suited the Frangibolt design. This section discusses the initial Frangi-HDRM (i.e., Frangiboltbased HDRM) development through engineering trials, qualification campaigns, and into flight  manufacture and in-orbit deployment of multiple modules. The development was challenging at times; an early release trial on some ground support equipment resulted in a non-fracture of the bolt, however,  these problems were overcome to fulfill the successful design qualification and release of multiple devices  in-orbit. Clock spring Deployed  telemetry  cam and  microswitchSolar array bracket  interface Craft bracket  interfaceEnd stop  limits Latching interface  and latch 424