Abstract: A scalable signal processing system is disclosed that processes digitized spectrum received from a constellation of satellites (or other sources), extracting multiple digital signals from multiple sources through multiple acquisition sites that is virtualized with high availability. A system of one or more antennas can receive a range of frequencies of raw spectrum covering multiple visible orbit planes, where a single antenna can receive signals from multiple satellite concurrently. This can be particularly useful when establishing a constellation satellites, where a number of satellites can be grouped together within an antenna's field of view. A group of digitizers receive the signals from the antennas and creates raw samples to form a spectrum sample pool. The spectrum sample pool is stored in a raw frame archive, where the digitizers and raw frame archive can be co-located and can also be co-located with one or more of the antennas.

Abstract: Described herein is a satellite communications system that includes: two or more satellites using laser communications, and a communications relay aircraft adapted for flying at altitudes above clouds. The communications relay aircraft includes: a laser communications module to communicate with the satellite using laser communication and a Radio Frequency (RF) communications module to communicate with RF equipment at or near ground level using cloud-penetrating RF communications. The RF communications module is configured to take data received as laser communication and generate a corresponding RF transmission containing the data. The laser communications module is configured to take data received as RF communication and to generate a corresponding laser transmission containing the data.

Abstract: Beamforming channels of a satellite are calibrated using a low power, spread spectrum calibration signal. The power of the calibration signal is below the noise level of a user signal in an active channel, allowing channels to be calibrated while active. When calibrating the transmit side circuitry, a two-stage calibration can be used, first calibrating the output hybrid matrix, then calibrating the whole of the transmit side. To improve performance, the dwell time spend calibrating a channel can be based on the power of the user signal in the channel. A transmit probe can be used to inject a calibration signal into the receive antennae and a receive probe can be used to extract the calibration signal from the transmit antennae. To reduce frequency of calibrations, the calibrations can be based on path-to-path differences. These techniques are also applied to multiport amplifiers (MPAs).

Abstract: A spacecraft includes a plurality of deployable module elements, at least one of the deployable module elements including a robotic manipulator, the spacecraft being reconfigurable from a launch configuration to an on-orbit configuration. In the launch configuration, the deployable module elements are disposed in a launch vehicle in a first arrangement. In the on-orbit configuration, the deployable module elements are disposed in a second configuration. The spacecraft is self-assembled by the robotic manipulator reconfiguring the spacecraft from the launch configuration, through a transition configuration, to the on-orbit configuration. The deployable module elements may be in a stacked arrangement in the launch configuration and may be in a side-by-side arrangement in the on-orbit configuration.

Abstract: A spacecraft includes payload equipment and bus equipment, each of the payload equipment and the bus equipment being coupled with a secondary structure arrangement. The spacecraft is configured to structurally interface with a launch vehicle upper stage only by way of an exoskeletal launch support structure (ELSS) that provides a first load path between a launch vehicle upper stage and the secondary structure arrangement. The spacecraft is configured to deploy from the launch vehicle upper stage by separating from the ELSS. In some implementations, the first load path is the only load path between the spacecraft and the launch vehicle upper stage, and substantially all of the ELSS remains with the launch vehicle upper stage. Because the secondary structure may be substantially less massy then the ELSS, subsequent orbit raising maneuvers may then be executed with a spacecraft having a reduced dry mass.

Abstract: In an on-orbit configuration, a spacecraft having a center of mass and a pitch axis passing through the center of mass includes a main body, a first solar array, and a first thruster is operable in a geostationary orbit with the first solar array deployed, proximate to a first north or south surface of the main body, such that a rotational axis of the deployed first solar array is substantially parallel to the pitch axis. The first thruster is disposed proximate to a second north or south surface of the main body, the first thruster having a thrust vector that is approximately coaligned with the pitch axis, the second surface being opposite to the first north or south surface. No solar array is proximate to the second surface.

Abstract: Techniques for articulating a sunshield to shade portions of a spacecraft are disclosed. In one aspect, a spacecraft includes a body and an articulable sunshield. The spacecraft is configured to operate in an orbital plane, such that the spacecraft 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 passing through the spacecraft coordinate system origin, and a roll axis orthogonal to the pitch axis and the yaw axis and passing through the spacecraft coordinate system origin. The sunshield is configured to rotate about an axis substantially parallel to the pitch axis such that a selected location of an exterior portion of the body is shaded from the Sun by a surface of the sunshield irrespective of seasonal and diurnal variations in orientation of the spacecraft with respect to the Sun.

Abstract: 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.

Abstract: A satellite communication system comprises one or more non-geostationary satellites. Each satellite is configured to provide a plurality of spot beams. The polarizations of the spot beams are laid out on each satellite so that terminals have a constant polarization as the field of regard traverses the terminal location.

Abstract: A first spacecraft and a second spacecraft are configured to be disposed together, in a launch configuration, for launch by a single launch vehicle. In the launch configuration, the first spacecraft is mechanically coupled with a primary payload adapter of the launch vehicle, and the second spacecraft is mechanically coupled with the first spacecraft by way of an inter-spacecraft coupling arrangement. The spacecraft are configured to be deployed, following injection into a first orbit by the launch vehicle, by separating the first spacecraft from the primary payload adapter while the second spacecraft is mechanically coupled with the first spacecraft. A first onboard propulsion subsystem of the first spacecraft includes one or more thrusters configured to execute an orbit transfer maneuver from the first orbit to a second orbit. A propellant line arrangement detachably couples the first onboard propulsion subsystem with a second propellant storage arrangement on the second spacecraft.

Abstract: This disclosure provides systems, methods and apparatus for a flexible radio frequency (RF) converter. In one aspect, a subsystem of a spacecraft can include a flexible RF converter having a printed circuit board (PCB), and a synthesizer and controller disposed on the PCB. The PCB can be placed within a chassis. The controller can communicate with the synthesizer to adjust a frequency of a synthesizer signal generated by the synthesizer, which is used to adjust a frequency of an input RF signal.

Abstract: A spacecraft includes a feed array that includes a plurality of unitary modules, each module including a respective plurality of radio frequency (RF) waveguides structurally coupled together with at least one connecting feature. For each unitary module, the at least one connecting feature and a wall structure defining the respective plurality of waveguides are co-fabricated using an additive manufacturing process. The plurality of unitary modules includes a first unitary module including a first plurality of waveguides and a second unitary module including a second plurality of waveguides. The first unitary module and the second unitary module are configured to be integrated together so as to electrically couple at least one of the first plurality of waveguides with at least one of the second plurality of waveguides.

Abstract: A spacecraft operating in a low earth orbit having an altitude in the range of 160 to 800 km has a main body that includes heat dissipating electrical equipment and an earth-facing payload. Control surfaces on the spacecraft are articulated so as to: provide three-axis attitude control to the spacecraft main body using aerodynamic drag effects, such that the earth-facing payload is maintained in a selected orientation with respect to the earth; and control one or both of orbit altitude and period by articulating the control surfaces so as to regulate an amount of aerodynamic drag. The control surfaces include a first control surface disposed, in an on-orbit configuration, on a boom, the boom being mechanically coupled with the main body, and with one or both of a solar array electrically coupled with the electrical equipment and a thermal radiating array thermally coupled with the electrical equipment.

Abstract: A multi-beam antenna (MBA) system for a spacecraft, the MBA system including a reflector and a feed array of radiating feed elements configured as a phased array and illuminating the reflector. The feed array includes a plurality of interchangeable modules. Each of the plurality of interchangeable modules includes a distal mounting panel and a proximal mounting panel, and at least six feed array elements. Each feed array element is electrically coupled with a respective amplifier and mechanically coupled with an exterior surface of the distal mounting panel. The respective amplifiers are thermally coupled with the proximal mounting panel and are mechanically coupled to an interior surface of the distal mounting panel and an exterior surface of the proximal mounting panel. An interior surface of the proximal mounting panel of each interchangeable module is mechanically and thermally coupled with a back plate.

Abstract: A constellation of Earth-orbiting spacecraft, the constellation having an orbital maneuver lifetime life (OML), includes a first spacecraft disposed in a first orbit and a second spacecraft disposed in a second orbit, each of orbit being substantially circular with a radius of approximately 42,164 km and having a respective inclination with respect to the equator specified within a range of 10° to 20°. The first orbit has, at beginning of life (BOL), a first right ascension of ascending node (BOL-RAAN1) and the second orbit has, at BOL, a second RAAN (BOL-RAAN2) the BOL-RAAN1 and the BOL-RAAN2 being separated by a first angular separation ?-RAAN1. A first stationkeeping delta-V (?V1) applied over the OML to the first spacecraft, in combination with a second delta-V (?V2) applied over the OML to the second spacecraft, maintains the ?-RAAN1 approximately constant and an actual inclination within specification, and ?V1 approximately equals ?V2.

Abstract: An amplifier system includes an input network having a plurality of input ports; an output network having a plurality of output ports; a plurality of amplification units coupled between the input network and the output network, the plurality of amplification units configured to amplify signals from the plurality of input ports; and a calibration unit coupled between the plurality of amplification units and the output network to calibrate amplified signals from the plurality of amplification units.

Abstract: A spacecraft includes a spacecraft main body, including payload equipment disposed on an inner panel surface of at least one radiator panel, the at least one radiator panel having an exterior-facing outer panel surface. The spacecraft includes at least one electric thruster disposed proximate to an aft facing panel of the spacecraft main body. The at least one electric thruster is thermally coupled with the at least one radiator panel by way of a thermally conductive path.

Abstract: Techniques for deploying a plurality of smallsats from a common launch vehicle are disclosed where a structural arrangement provides a load path between an upper stage of the launch and the plurality of spacecraft. Each spacecraft is mechanically coupled with the launch vehicle upper stage only by the structural arrangement. The structural arrangement includes at least one trunk member that is approximately aligned with the longitudinal axis of the launch vehicle upper stage, a plurality of branch members, each branch member being attached to the trunk member and having at least a first end portion that is substantially outboard from the longitudinal axis; and a plurality of mechanical linkages, each linkage coupled at a first end with a first respective spacecraft and coupled at a second end with one of the plurality of branch members, the trunk member or a second respective spacecraft.

Abstract: Techniques for orienting an earth-orbiting spacecraft include determining, using a star tracker on board the spacecraft, a first vector aligned between an ecliptic pole of the earth and the spacecraft, adjusting attitude of the spacecraft so as to align a first axis of the spacecraft with the first vector, and rotating the spacecraft about the first axis until presence of the sun is registered. Rotation rates may be subsequently reduced, such that the sun remains within a field of view of the sun sensor or of a solar array of the spacecraft.

Abstract: An amplifier system includes an input network having a plurality of input ports; an output network having a plurality of output ports; a plurality of amplification units coupled between the input network and the output network, the plurality of amplification units configured to amplify signals from the plurality of input ports; and a calibration unit coupled between the plurality of amplification units and the output network to calibrate amplified signals from the plurality of amplification units.