Abstract: A system for harvesting solar energy on a spacecraft includes a stiff substrate layer and a working layer disposed on the substrate layer to provide at least one of a photovoltaic or a reflective function. In a first operational state, the substrate layer is arranged as a tape spring to store potential energy which causes the substrate layer to uncoil and provide, in a second operational state, a photovoltaic module and/or a solar concentrator.

Abstract: An orbital satellite has a bus formed of a bus module and a payload module. Both the bus module and the payload modules may include a pair of panels integrally formed at an angle of for example 90° with respect to each other. Components and electronics supporting the satellite systems may be mounted on the panels of the bus and payload modules. Heat generated by the components and electronics are conducted between panels of the bus module and/or payload module. Sufficient heat transfer between panels occurs as a result of their being integrally formed with each other.

Abstract: A solar array structure, such as for a spacecraft, uses thin solar array panels that, when in a stowed configuration, are stiffened by being bent or curved in one direction to be shaped like a section of a cylinder and placed within a rigid structural frame. As a curved solar panel is not as efficient as a flat panel directly facing the sun, the solar array panels are curved in their stowed configuration for launch only, but flatten after deployment by use of a partially flexible structural frame, where a rectangular frame is made of two opposing rigid sides and two opposing flexible sides, with a thin flexible solar panel attached to rigid sides only. The rigid sides are compressed during stowage to curve the panel before hold-down tensioning. The structure and panels return to their flat free state configuration after release.

Abstract: A solar array structure for a spacecraft is based on a modular approach, allowing for arrays to be designed, and designed to be modified, and manufactured in reduced time and with reduced cost. The embodiments for the solar array are formed of multiple copies of a “bay” of a multiple strings of solar array cells mounted on semi-rigid face-sheet structural elements. The bays are then placed into frame structures made of tubes connected by nodes to provide an easily scalable, configurable, and producible solar array wing structure. This allows for rapid turnaround of program specific designs and proposal iterations that is quickly adaptable to new/future PhotoVoltaic (PV) technologies and that can create uniquely shaped (i.e., not rectangular) arrays, allowing for mass production with simple mass producible building blocks.

Abstract: An apparatus includes an integral, additively manufactured, actuation device having a rigid portion comprising a shaped structural member and a flexible portion comprising a helical torsion spring. In a spacecraft application, a spacecraft appendage may be coupled with a deployment mechanism, the deployment mechanism including at least one integral, additively manufactured, actuation device having a rigid portion comprising a shaped structural member and a flexible portion comprising a helical torsion spring.

Abstract: An attitude control system for a satellite is presented that can determine on time values for attitude control thrusters without use of attitude rate sensors, such as those based on gyros. The attitude control systems uses the attitude values from a star tracker to determine both attitude adjustment values and attitude adjustment rate values directly from the star tracker values, where the processing is performed using quaternions. From the attitude adjustment values and attitude adjustment rate values, a set of thruster on time values are determined.

Abstract: A solar array structure for a spacecraft is based on a modular approach, allowing for arrays to be designed, and designed to be modified, and manufactured in reduced time and with reduced cost. The embodiments for the solar array are formed of multiple copies of a “bay” of a multiple strings of solar array cells mounted on semi-rigid face-sheet structural elements. The bays are then placed into frame structures made of tubes connected by nodes to provide an easily scalable, configurable, and producible solar array wing structure. This allows for rapid turnaround of program specific designs and proposal iterations that is quickly adaptable to new/future PhotoVoltaic (PV) technologies and that can create uniquely shaped (i.e., not rectangular) arrays, allowing for mass production with simple mass producible building blocks.

Abstract: A method of roll steering of a spacecraft to align an aspect of the spacecraft, such as the surface of solar arrays carried by the spacecraft, to the sun, is described. The roll steering occurs only when the sun is at an angle (?) relative to the orbital plane of the spacecraft and when the spacecraft is not eclipsed by a body it is orbiting. This dayside-only roll steering of the spacecraft increases the power efficiency of the spacecraft. A spacecraft may include a controller which causes an attitude control subsystem to steer the spacecraft about a roll axis to position the surface of the solar array such that an axis normal to the surface of the solar array is aligned with the direction to a sun when the sun is visible to the spacecraft, and maintain a fixed orientation of the spacecraft about the roll axis when the sun is not visible to the spacecraft.

Abstract: A satellite observation system and method of deploying a satellite system are disclosed. The system includes a plurality of observation satellites comprising one or more sensors, each of the plurality of observation satellites configured with at least a solar array and a mechanical stabilization element. Each of the plurality of observation satellites is constructed without positioning components. The plurality of observation satellites is positioned in a dawn/dusk sun-synchronous orbital plane about a celestial body such that the one or more observation sensors are oriented toward the celestial body.

Abstract: Technology is disclosed herein for orbit raising of multiple spacecraft launched with a single launch vehicle. Two or more spacecraft are configured in a stacked launch configuration in which a lower spacecraft is mechanically coupled with a payload adapter of a launch vehicle with one or more upper spacecraft above the lower spacecraft. Propellant that is stored in the lower spacecraft during launch is transferred to an upper spacecraft in the stack after launch. The propellent may be used by the upper spacecraft for an orbit raising maneuver that raises the orbit of at least the upper spacecraft from a first orbit to a second orbit. Storing the propellant in the lower spacecraft lowers the center of mass of the stack during launch. Lowering the center of mass reduces the structural bending moment of the stack during launch, which allows a greater total launch mass.

Abstract: A voltage measurement apparatus for a spacecraft thruster includes a thruster interface having a platform and at least one spacer element configured to attach the platform to the thruster. At least one probe is mounted to the thruster interface with the probe configured to engage an anode assembly of the thruster when the thruster interface is attached to the thruster. The apparatus also includes voltage metering circuitry coupled to the at least one probe, the voltage metering circuitry configured to be powered by the spacecraft thruster when the spacecraft thruster is powered on.

Abstract: Technology is disclosed herein for reverse current protection for a photovoltaic module string. An apparatus has a switch connected between each respective PV module string and a power bus. A control circuit closes a set of the switches to connect a set of the PV module strings to the power bus to transfer power from the set of the photovoltaic PV module strings to the power bus. The control circuit determines whether a reverse current flows in the direction from the power bus to any of the PV module strings. The control circuit maintains the switch associated with a particular PV module string in an open state to prevent reverse current from flowing in the particular PV module string responsive to a determination that a reverse current flows in the particular PV module string when the particular PV module string is connected to the power bus.

Abstract: Technology is disclosed herein for a power control and distribution unit (PCDU) of a spacecraft that has a microcontroller to control battery charging from solar arrays. Using a microcontroller within the PCDU reduces the complexity of the PCDU. The microcontroller may be programmable and reprogrammable, which allows the charging of the battery to be adapted to various conditions. For example, the microcontroller can be programmed in accordance with the mission to optimize battery charging for that mission.

Abstract: Technology is disclosed herein for using an antenna reflector to regulate a spacecraft temperature during orbit raising. When in a launch configuration, the antenna reflector may be stowed in a fairing of a launch vehicle. After the spacecraft is deployed from the launch vehicle and prior to orbit raising, the antenna reflector is moved from the launch configuration to an orbit raising configuration in which the antenna reflector is used to regulate the spacecraft temperature. The antenna reflector may be proximate a thermal radiator panel of the spacecraft when in the launch configuration. The antenna reflector may be positioned such that sunlight will reflect off the antenna reflector onto the thermal radiator panel, thereby warming the spacecraft. After orbit raising, the antenna reflector is moved from the orbit raising configuration to an operational configuration in which a boresight of the antenna reflector may be directed toward nadir.

Abstract: Techniques for minimizing diurnal temperature variation of a radiator of a spacecraft are disclosed. In one aspect, a spacecraft includes a body, a radiator panel, and a heat dissipating unit thermally coupled with the radiator panel. The spacecraft is configured to operate in an orbital plane, and has a yaw axis within the orbital plane and directed from a spacecraft coordinate system origin toward nadir, a pitch axis orthogonal to the orbital plane, and a roll axis orthogonal to the pitch axis and the yaw axis. The radiator panel includes a surface area external to a body of the spacecraft, a first portion of the surface area facing a first direction that is substantially parallel to the roll axis, and a second portion of the surface area facing a second direction that has a substantial component parallel to the yaw axis.

Abstract: A universal test port is connected to the different functional sub-systems of a spacecraft, allowing the sub-systems to be tested from a single location of an assembled spacecraft. The universal test port is mounted on an external surface of the spacecraft and configured to connect to the different functional sub-systems (such as power, propulsion, and command and data handling, for example) of the assembled spacecraft, allowing for the streamlining of testing operations by electrical ground system equipment during assembly, integration, and test (AIT) operations and reducing the risk of collateral damage to spacecraft hardware during testing in AIT.

Abstract: Technology is disclosed herein for reverse current protection for a photovoltaic module string. An apparatus has a switch connected between each respective PV module string and a power bus. A control circuit closes a set of the switches to connect a set of the PV module strings to the power bus to transfer power from the set of the photovoltaic PV module strings to the power bus. The control circuit determines whether a reverse current flows in the direction from the power bus to any of the PV module strings. The control circuit maintains the switch associated with a particular PV module string in an open state to prevent reverse current from flowing in the particular PV module string responsive to a determination that a reverse current flows in the particular PV module string when the particular PV module string is connected to the power bus.