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.

Abstract: An example of an apparatus includes an inlet to connect to a propellant source a pressure regulator connected to the inlet to reduce propellant pressure from a first pressure at the inlet to a second pressure. The apparatus includes a manifold connected to the pressure regulator to receive propellant from the pressure regulator at the second pressure and includes a plurality of manifold outlets. The apparatus further includes a plurality of gas lines, each gas line extending from a corresponding manifold outlet for connection to a corresponding satellite propellant tank.

Abstract: 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.

Abstract: A solar array structure for a spacecraft includes one or a pair of flexible blanket or other foldable solar arrays 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.

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: A solar array structure for a spacecraft includes one or a pair of flexible blanket or other foldable solar arrays 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.

Abstract: Technology is disclosed for a dispenserless multi-satellite launch configuration in which multiple satellites are interconnected to form a composite beam structure that provides stability independently of the launch vehicle. When in the launch configuration, the satellites are formed into a bundle, where each satellite connects by one or more simple connector along the edges of its inner facing vertical side to the satellite adjacent on each side. This composite beam structure provides a stable launch configuration independently of the launch vehicle. Each of the satellites also has one or more connectors along the bottom edge of the inner facing vertical side allowing the bundle to be attached to a ring type launch vehicle interface. Once launched, the satellites can be dispensed by releasing the connector.

Abstract: A satellite includes a first radiator panel, a second radiator panel, a space defined between the first radiator panel and the second radiator panel, and one or more first heat-generating components located in the space. Each of the first heat-generating components is attached to at least one of the first or second radiator panels. The satellite further includes a third radiator panel extending from the space and one or more second heat-generating components located in the space, each of the second heat-generating components is attached to the third radiator panel.