Abstract: Technology is disclosed herein for dispensing stacked spacecraft. A stack of spacecraft may be joined together in an accordion configuration by dis-engageable links. A dis-engageable link may join two adjacent spacecraft at one edge of the spacecraft. Some of the links are on one side of the stack with other links on an opposite side of the stack to provide the accordion configuration. After the tie-down mechanism releases the stack of spacecraft from a launch adaptor the stack unfolds. Initially, the dis-engageable links continue to hold the spacecraft together in the accordion configuration with the angle between each pair of adjacent spacecraft increasing. After a pair of adjacent spacecraft have unfolded a sufficient amount to prevent collision, the dis-engageable links release one of the spacecraft to thereby dispense the spacecraft.

Abstract: A distributed computer system for a spacecraft is disclosed. The system has multiple computer nodes, each controlling a different aspect of a mission of the spacecraft. Each node includes a control circuit(s) that controls a set of components, a router processor, and a programmable processor. The programmable processor of each respective computer node issue commands to the control circuit(s) of the respective computer node to carry out an aspect of the mission associated with the respective computer node. Upon failure of the programmable processor in a particular computer node, a healthy programmable processor send commands to the router processor in the particular computer node The router processor of the particular computer node routes the commands received from the remote programmable processor to the control circuit(s) in the particular computer node to control the set of components to carry out the aspect of the mission associated with particular computer node.

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 distributed computer system for a spacecraft is disclosed. The system has multiple computer nodes, each controlling a different aspect of a mission of the spacecraft. Each node includes a control circuit(s) that controls a set of components, a router processor, and a programmable processor. The programmable processor of each respective computer node issue commands to the control circuit(s) of the respective computer node to carry out an aspect of the mission associated with the respective computer node. Upon failure of the programmable processor in a particular computer node, a healthy programmable processor send commands to the router processor in the particular computer node The router processor of the particular computer node routes the commands received from the remote programmable processor to the control circuit(s) in the particular computer node to control the set of components to carry out the aspect of the mission associated with particular computer node.

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: An example of an apparatus includes a spacecraft body and a solar array that is attached to the spacecraft body. In addition, a solar array spring element is configured to provide a force between the solar array and one or more components of a neighboring spacecraft in a stack of spacecraft.

Abstract: Methods and structures are presented for implementing an automatic target recognition system as a convolutional neural network (CNN) in a satellite or other environment with constrained resources, such as limited memory capacity and limited processing capability. For example, this allows for the automatic target recognition to be implemented on a field programmable gate array (FPGA). Image data is split into subsets of contiguous pixels, with the subsets processed in parallel in a CNN of a corresponding processing node using quantized weight values that are determined in a training process that accounts for the constraints of the automatic target recognition system. The results of the automatic target recognition process is based on the combined output of the processing nodes.

Abstract: Technology 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: An example apparatus includes a first 3D printer head configured to form a spiral structure around a hub and a second 3D printer head configured to form a boom extending between the second 3D printer head and the hub. The apparatus further includes one or more actuators coupled to the first 3D printer head and the second 3D printer head to control a distance between the first 3D printer head and the second 3D printer head.

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 passively damped mechanical system is disclosed, for example for use in aerospace applications where vibration can adversely affect navigational and operational instruments. In one example, the passively damped mechanical system includes an end fitting of a strut used to connect a structural element to a payload. The end fitting may include outer and inner cylindrical hubs, with a space between the outer and inner cylindrical hub at least partially filled with a viscoelastic material. In a further example, the passively damped mechanical system includes legs used to connect a structural element to a bracket configured to support a payload. Each leg may include a hollow interior having a lattice structure to add strength and a viscoelastic material to provide passive damping.

Abstract: An orbital satellite has a pair of multi-axis booms including both thrusters for course/attitude adjustment and an end effector for grappling payloads and manipulating other tools and objects. The satellite may launch with a primary payload affixed to a bus and one or more secondary payloads affixed to an ESPA ring. Once in orbit, the end effector may be used to grapple the primary and/or secondary payloads and rearrange them on the bus. In further aspects, the end effector may be used to make bus repairs or take measurements, or hold tools that are used to make bus repairs or take measurements.

Abstract: A satellite includes a first radiator panel with first heat-generating components attached to its surface and a second radiator panel with second heat-generating components attached to its surface. One or more actuators are configured to deploy the first and second radiator panels from a compact configuration in which the first and second radiator panels are overlapping to a deployed configuration in which the first and second radiator panels are non-overlapping.

Abstract: 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.

Abstract: Technology is disclosed herein for a payload dispensing hinge assembly. The payload dispensing hinge assembly has a first hinge-half and a second hinge-half that are joined by a hinge pin. The first hinge-half may be connected to a payload that is to be dispensed at a target angle. The second hinge-half may be connected to a payload base. The first hinge-half may have a first mounting bracket and a rotatable arm that are shaped to form interlocks that serve to dis-engageably link these two components. A biasing mechanism rotates the rotatable arm and hence the first mounting bracket and payload about a hinge line. The second hinge-half has a hinge stop that stops the rotation of the rotatable arm at a target angle, whereby the first bracket dis-engages from the rotatable arm to dispense the payload at the target angle.

Abstract: Technology is disclosed herein for dispensing stacked spacecraft. A stack of spacecraft may be joined together in an accordion configuration by dis-engageable links. A dis-engageable link may join two adjacent spacecraft at one edge of the spacecraft. Some of the links are on one side of the stack with other links on an opposite side of the stack to provide the accordion configuration. After the tie-down mechanism releases the stack of spacecraft from a launch adaptor the stack unfolds. Initially, the dis-engageable links continue to hold the spacecraft together in the accordion configuration with the angle between each pair of adjacent spacecraft increasing. After a pair of adjacent spacecraft have unfolded a sufficient amount to prevent collision, the dis-engageable links release one of the spacecraft to thereby dispense the spacecraft.

Abstract: Technology is disclosed herein for a payload dispensing hinge assembly. The payload dispensing hinge assembly has a first hinge-half and a second hinge-half that are joined by a hinge pin. The first hinge-half may be connected to a payload that is to be dispensed at a target angle. The second hinge-half may be connected to a payload base. The first hinge-half may have a first mounting bracket and a rotatable arm that are shaped to form interlocks that serve to dis-engageably link these two components. A biasing mechanism rotates the rotatable arm and hence the first mounting bracket and payload about a hinge line. The second hinge-half has a hinge stop that stops the rotation of the rotatable arm at a target angle, whereby the first bracket dis-engages from the rotatable arm to dispense the payload at the target angle.

Abstract: A passively damped mechanical system is disclosed, for example for use in aerospace applications where vibration can adversely affect navigational and operational instruments. In one example, the passively damped mechanical system includes an end fitting of a strut used to connect a structural element to a payload. The end fitting may include outer and inner cylindrical hubs, with a space between the outer and inner cylindrical hub at least partially filled with a viscoelastic material. In a further example, the passively damped mechanical system includes legs used to connect a structural element to a bracket configured to support a payload. Each leg may include a hollow interior having a lattice structure to add strength and a viscoelastic material to provide passive damping.

Abstract: A satellite propellant tank includes a tank body and a dome attached to the tank body to enclose an interior volume for propellant storage. One or more cavities are formed in the dome. One or more propellant control components are located in the one or more cavities formed in the dome.

Abstract: An example of an apparatus includes a spacecraft body and a solar array that is attached to the spacecraft body. In addition, a solar array spring element is configured to provide a force between the solar array and one or more components of a neighboring spacecraft in a stack of spacecraft.