Abstract: A process for designing spacecraft structural elements (20, 30) that increases spacecraft structure intrinsic damping to relax stiffness design constraints that are necessary for precision pointing requirements. The process includes specifically designing the spacecraft structural elements (20, 30) to have a stiffness that is intrinsically not suitable to meet mission pointing performance requirements in order to reduce weight and volume. To overcome this deficiency, the structural elements (20, 30) are equipped with strain energy control elements (44) that sense strain in the structural elements (20, 30) from on-board and external disturbances, and provide actuation of the structural elements (20, 30) to counteract the sensed strain. The strain energy control elements (44) can be any suitable control element that senses strain and actuates the structural element (20, 30), such as piezoelectric electric or electrostrictive control elements.

Abstract: The power to weight ratio obtained with a spacecraft's solar array (3) is enhanced by a factor of at least two to five through use of support structure for individual solar panels or solar arrays (31a-17a) containing a curved outwardly bowed surface as deployed that packs essentially flat for storage. Defining a D-shaped wing (15a) in cross-section as deployed, the support structure (4 & 6) for a string of solar cells (10, FIG. 3) provides greater inertia and thereby greater rigidity than prior designs. One member (4) to the support structure is relatively flexible. One-hundred and eighty degree strain energy hinges (19) carried by the other support member (6) outwardly bows that flexible to define a curved sector when the support structure is released from the stowed condition in which the support structure is held relatively thin and flat.

Abstract: A technique for measuring distortion in a structure of interest, such as a spacecraft antenna reflector (18), and optionally compensating for the distortion. A first set of targets (22) on the structure (18) is scanned by an attitude transfer system (24) to measure the angular location and range of each target relative to a reference point on another structure (12) having a frame of reference. The orientation of the structure of interest is them determined from the measured locations of the targets. A second set of targets (60 or 82) on the structure of interest is scanned by a figure sensing module (26) located at a reference point on the structure itself. From measured angular locations and ranges of the second set of targets, any shape distortion in the structure of interest can be determined, and distortion may be corrected with the use of actuators (98).

Abstract: The power to weight ratio obtained with a spacecraft's solar array (3) is enhanced by a factor of at least two to five through use of support structure for individual solar panels or solar arrays (31a-17a) containing a curved outwardly bowed surface as deployed that packs essentially flat for storage. Defining a D-shaped wing (15a) in cross-section as deployed, the support structure (4 & 6) for a string of solar cells (10, FIG. 3) provides greater inertia and thereby greater rigidity than prior designs. One member (4) to the support structure is relatively flexible. One-hundred and eighty degree strain energy hinges (19) carried by the other support member (6) outwardly bows that flexible to define a curved sector when the support structure is released from the stowed condition in which the support structure is held relatively thin and flat.

Abstract: A spacecraft structure using functionally independent modules assembled around a lightweight core structure to provide a vehicle that is lighter, uses less volume, and is easier to design, manufacture and test than a conventional spacecraft. In the disclosed embodiments, the modules are formed on generally flat panels, which serve as thermal radiators. The modules extend radially from the core structure and are attached to the core structure either in coplanar rows that extend axially along the core structure or in a coplanar set that extends circumferentially around the core structure. Interconnection between modules is achieved through a backbone interface, through which the modules are connected to the core structure. A large number of variant configurations may be implemented using the modular approach, by selecting a core, components and modules of number and size to meet mission requirements.

Abstract: A technique for reducing the acoustic response of a spacecraft electronics equipment module that is structurally and thermally independent of a core spacecraft structure to which it can be mounted. The module takes the form of a thermal radiation panel on which electronic components are directly mounted. The panel is adjusted in stiffness to reduce its dynamic vibration response to acoustics and launch transients. These adjustments are facilitated by the external mounting of the panel to the core structure, preferably using a statically determinate mount that renders the module even less susceptible to vibration transferred from the core structure. External mounting removes the panel from the primary structure load path, which provides the freedom to adjust the stiffness as desired.