a deployment drive actuator, deployment latch and a load-bearing structural member when deployed. This  combination of attributes could potentially provide significant mass, complexity and cost savings for  spacecraft deployable structures.   The HSC field has grown significantly since the early 2000s and is well defined in the seminal paper by  Murphey et al.13 One recent flight success involving HSC materials was the Air Force Research Laboratory  Roll-Out Solar Array that flew on the International Space Station in 2017.14 The use of HSC materials for  precision hinges was originally shown in the literature by Domber et al.12 As indicated in Table 1, their work  found submicron-level axial and transverse deployment precision for one HSC hinge design and higher  precision error for two other hinges made from different composite layups. Since their work, many new HSC  materials have been developed and high precision position sensors have become available enabling the  measurement of more degrees-of-freedom. These new developments motivated the current work. In this  paper, deployment precision performance in four degrees-of-freedom is presented for hinges of two  different geometries and six different composite layups.   To make HSC hinges, two opposing tape sections are bonded together at the ends using end fittings. The  hinges can be folding by locally buckling the tapes to flatten them at their midspan, then bending the tapes  in the flattened region to fold the hinge. An HSC hinge is shown deployed and stowed in Figure 2.   Figure 2. Hinge deployed (left) and stowed (right)  With no specific system in mind for the present study, the two HSC hinge geometries chosen for testing  were selected based upon their applicability to small satellite deployable structures, available composite  tooling, and available oven sizes. For CubeSat-scale small satellites, the maximum usable width along one  side of the spacecraft is 8.5 cm. Using this limit as the flattened width of the tapes, only two available  cylindrical composite mandrels were available to make hinges that could fold without damage: 15.2-cm (6inch) and 10.2-cm (4-inch)-diameter mandrels. The free-edge length of 30 cm was chosen based on finite  element analysis to avoid damage inducing strain concentrations near the end fittings when folded and so  that the final hinge bonding can be done in an available oven.  The materials chosen for testing were based on materials available at the Air Force Research Laboratory  (AFRL) Space Vehicles Directorate and the layups chosen were based on AFRL staff recommendations  and previous publications.15, 16  The materials and layups chosen for this work are listed in Table 2. At least  two tapes of each layup and diameter were made at the AFRL and provided to MITLL. At MITLL, the tapes  were then trimmed to size and two tapes were bonded to end brackets to make hinge specimens. The  bracket pieces for the fixed end were made from Invar to minimize thermally induced shape change. The  bracket pieces for the free end were made from aluminum and significantly light weighted to minimize  gravity-loading effects at the tip. The total mass of the light weighted bracket assembly is 173 grams. The  base and tip bracket assemblies were bonded to the tapes in one operation using Hysol EA 9394 epoxy  with 180-µm glass beads to enforce a minimum bondline thickness. The tip and base bracket assemblies  included alignment features to interface with tooling during the bonding process to set the distance and  NASA/CP—2018-219887 419