Maxar Space LLC

A connector loading device may comprise a receptacle, a plurality of light sources, and a controller. The receptacle may comprise a first plurality of holes. The first plurality of holes are configured to respectively line up with a second plurality of holes in a connector when the connector is inserted in the receptacle. Each one of the first plurality of holes comprises a respective and corresponding one of a plurality of unique indexes. The plurality of light sources respectively correspond to the first plurality of holes. Each one of the plurality of light sources, when initiated, is configured to illuminate light from a bottom of its respective one of the first plurality of holes through a top of its respective one of the first plurality of holes. The controller is configured to receive an input and initiate a one of the plurality of light sources corresponding to the index.

A solid state power amplifier uses a Doherty power amplifier that can be implemented as a monolithic microwave integrated circuit. By adjusting the DC bias of the amplifying stages in each branch of the Doherty amplifier, the output power, linearity, and DC power can be adjusted to provide a specified output, where the specification for the output can include the maintaining of desired DC power and linearity. The Doherty power amplifier can be used in a satellite payload or other application utilizing solid state power amplifiers, while providing the proper amount of RF output power and DC power. A single amplifier can have its bias levels adjusted for different output levels, helping to minimize the number of designs that are required for a given satellite payload, reducing the variety of parts in a satellite payload.

A radiator structure for a satellite is provided. A first radiator panel adapted to be positioned on a first side of a central body, and a second radiator panel adapted to be positioned on a second side of the central body. Other implementations include a third radiator panel positioned on a third side of the central body. The apparatus also includes at least one heat pipe embedded between a first face and a second face of each radiator panel and extending from the first radiator panel through the first radiator panel and through the second radiator panel. The heat pipe structurally supports the first radiator panel and the second radiator panel relative to the intermediate radiator panel. A method of manufacturing a radiator structure is also provided.

To facilitate on-orbit servicing, such as for a refueling operation, techniques are presented for a servicing satellite to cut through the multi-layer insulation blanket of a client satellite to provide access to the client satellite without releasing unacceptable quantities of foreign object debris from the multi-layer insulation. The serving satellite includes a sealing tool, such as a pair of heater rollers, that apply pressure and heat to the insulating blanket to melt the inner layers and seal the outer layers together. The servicing satellite can then use a cutting tool to cut the sealed region and access the client satellite.

An electric propulsion simulator console (EPSC) which electronically simulates an electric propulsion assembly of a spacecraft as well as propulsion fuel control components and positioning components of the spacecraft. The EPSC simulates a spacecraft thruster electrical interface can test four thruster interfaces simultaneously and continuously. The simulator additionally facilitates the testing of spacecraft fault detection, isolation, and recovery by simulating failed magnet circuits, open anode paths, and flameout conditions. The EPSC includes an electrical propulsion unit load simulator adapted to receive propulsion unit control signals from a spacecraft under test and a spacecraft propulsion unit positioner simulator the simulator adapted to display a simulated state of three axes of movement for at least one propulsion unit positioner responsive to positioning signals received from the spacecraft under test. A propulsion unit fuel valve simulator is also provided and can display a simulated state of propulsion unit fuel valves responsive to control signals received from the spacecraft under test.

Technology is disclosed herein for preventing or at least significantly reducing shock to spacecraft such as satellites when releasing a hold-down rod assembly that clamps the spacecraft to, for example, a launch vehicle adaptor. The hold-down rod assembly has tension rods that may be pre-loaded at considerable tension in order to hold down a stack of spacecraft in a launch configuration. In an embodiment, pneumatic actuators are used to slowly release the tension in the tension rods. Therefore, shock to the spacecraft is prevented or at least significantly reduced.

A solar array structure for a spacecraft includes one or a pair of flexible blanket or other foldable solar arrays (such as flexible panels) and a deployable frame structure. The deployable frame structure includes a T-shaped yoke structure, a T-shaped end structure, and one or more rigid beams, the T-shaped yoke structure connectable to the spacecraft. When deployed, the frame structure tensions the flexible blanket solar array or arrays between the T-shaped yoke structure and the T-shaped end structure. When stowed, the flexible blanket solar array or arrays are folded in an accordion manner to form a stowed pack or packs between the cross-member arms of the T-shaped yoke structure and the T-shaped end structure, also stowed in its own Z-fold arrangement. The cross-member arms of the T-shaped end structure can include a solar array that can provide power before deployment while the flexible blanket solar array is stowed.

Technology is disclosed herein for deploying stacked spacecraft. When the spacecraft are stacked corresponding z-axis magnetic torque rods of the spacecraft will align with each other along a z-axis. Thus, collectively the stack of spacecraft have one or more sets of magnetic torque rods aligned with the z-axis. Just prior to deploying the spacecraft the one or more sets of magnetic torque rods are operated to hold the stack of spacecraft together. For example, the north magnetic pole of the magnetic torque rod in one spacecraft may face the south magnetic pole of the magnetic torque rod in an adjacent spacecraft. To deploy the top spacecraft, the polarity of the z-axis magnetic torque rod(s) in the top spacecraft is/are reversed. After the spacecraft is clear of the stack the magnetic torque rod(s) in the deployed spacecraft may be de-activated. Then another spacecraft may be deployed in a similar manner.

A Radio Frequency Identification (“RFID”) sensor harvests power from a RFID interrogation signal received from a RFID reader and harvests power from surrounding RF energy which is at a different frequency than the RF interrogation signal. The harvested power is used to operate the sensor and transmit data sensed by the sensor.

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.