Abstract: A flow distributor for a heat transfer device having a plurality of channels includes a sheath defining a plurality of distributor holes, each distributor hole configured to be in fluid communication with a respective channel inlet of each channel of the heat transfer device and an insert defining a plurality of fluid channels therein and a fluid inlet, each fluid channel in fluid communication with the fluid inlet. The insert is disposed within the sheath to seal the fluid channels with each fluid channel in fluid communication with a respective one of the distribution holes. The fluid inlet includes an inner inlet and an outer inlet radially outward from the inner inlet for mixing a fluid flow in the fluid inlet for evenly distributing fluid flow (e.g., a two phase flow) into the fluid channels of the insert and into each channel of the heat transfer device.

Abstract: A deployable membrane structure for an antenna comprises a membrane comprising a plurality of first regions of higher-stiffness material integrally connected via one or more second regions of lower-stiffness material, wherein the one or more second regions are formed from compliant material configured to permit the membrane to be folded into a collapsed configuration and subsequently unfolded into a deployed configuration, and are arranged so as to allow adjacent ones of the plurality of first regions to be folded so as to lie against one another. In some embodiments the membrane is formed of a composite material comprising a plurality of fibres in a compliant matrix, and the plurality of first regions comprise material with a higher fibre density than the one or more second regions. A deployable antenna comprising the deployable membrane structure is also disclosed.

Abstract: A spacecraft includes a propulsion system that includes one or more pressurant tanks configured to store an inert gas at a high pressure, one or more propellant tanks configured to store liquid propellant at an intermediate pressure, electric thrusters operable with the inert gas at a low pressure and pneumatically coupled with the one or more pressurant tanks by way of a first pressure regulator, and chemical thrusters operable with the liquid propellant. The inert gas is one or a mixture of two or more of xenon, argon and krypton. At least a portion of the liquid propellant is stored in at least one of the propellant tanks, the propellant tank including an ullage volume pneumatically coupled with at least one of the pressurant tanks by way of a second pressure regulator having an output set to the intermediate pressure and the ullage volume is pressurized by the inert gas.

Abstract: A spacecraft includes an additive manufacturing (A/M) subsystem and one or both of a thermal control arrangement and a contamination control arrangement. The A/M subsystem includes an A/M tool, feedstock and a workpiece and is configured to additively manufacture the workpiece using material from the feedstock. The thermal control arrangement is operable, in an on-orbit space environment characterized by near vacuum pressure and near zero-g force, to maintain temperature of at least one of the A/M tool, the feedstock, and the workpiece within respective specified ranges. The contamination control arrangement is operable, in the on-orbit space environment, to control outgassing of volatile organic compounds (VOCs).

Abstract: A spacecraft includes a power train that includes a solar array, an electric propulsion subsystem, and a power conversion module. The power conversion module receives power from the solar array at a voltage, Vi; and delivers power to the electric propulsion subsystem at a voltage, Vo. The spacecraft is configured to operate at a varying distance, D, from the sun within the range of Dmin to Dmax, Dmax being at least 1.3×Dmin. The solar array is configured to deliver power to the power conversion module at a voltage approximately equal to Vo when the spacecraft is proximate to Dmax.

Abstract: A spacecraft, reconfigurable from a launch configuration to an on-orbit configuration, includes a main body structure, a manipulator, a first deployable rigid reflector and an attachment arrangement, including at least one hold-down assembly (HDA). In the launch configuration, the HDA is in a fully engaged configuration such that the attachment arrangement mechanically attaches the first reflector with the spacecraft main body structure and prevents relative motion between the first reflector and the spacecraft main body. Reconfiguring the spacecraft from the launch configuration to the on-orbit configuration includes (i) actuating the HDA from the fully engaged configuration to a partially engaged configuration; (ii) grasping and moving the first reflector, with the manipulator, a distance in the first direction; and (iii) moving the first reflector from a first position proximate to the attachment arrangement to a second position proximate to a deployed position associated with the on-orbit configuration.

Abstract: An apparatus for launching, recovering, transporting, and storing vehicles is disclosed. The apparatus stabilizes the vehicle while it is in operation or inactive and has a frame connected to at least one stabilizer. In certain configurations, the apparatus stabilizes the vehicle without using the vehicle's onboard landing gear. The apparatus may also include at least one pad connected to the apparatus.

Abstract: A spacecraft is disclosed, comprising a deployable spacecraft body (110) comprising a plurality of sub-systems (321-324) for controlling operations of the spacecraft, and a plurality of panels (101, 102) and a plurality of hinges (112-115) each connecting adjacent ones of the plurality of panels, the hinges being arranged to permit the plurality of panels to be folded into a stowed configuration and unfolded into a deployed configuration, wherein the plurality of sub-systems are fixed to and supported by one or more of the plurality of panels. By forming the body of the spacecraft from a deployable structure, the overall size of the spacecraft can be significantly reduced in the stowed configuration.

Abstract: A first spacecraft includes a first fluid storage arrangement and a fluid flow metering arrangement including a holding tank coupled with the first fluid storage arrangement, a flow meter disposed proximate to the holding tank, and an active thermal control arrangement controlling the temperature of the flow meter and the holding tank. The first spacecraft is configured to service a second spacecraft, the second spacecraft including a second fluid storage arrangement, by transferring one or both of a propellant and a pressurant from the first fluid storage arrangement to the holding tank, and from the holding tank, through the flow meter, to the second fluid storage arrangement.

Abstract: The invention relates to a device (1) for use in collection and removal of slurry or similar from a sty, a sty comprising such a device and to a method of installing a device at a sty, such as a pig sty, wherein the device comprises an one-piece body (2) having one or more funnels (3) wherein each funnel comprises at least one upper opening (3a) having an upper horizontal area enclosed by an upper edge (3a?) and a lower opening (3b) having a lower horizontal area enclosed by a lower edge (3b?), wherein the upper horizontal area is larger than the lower horizontal area, and the inside (3?) of the funnel is thus extending from the upper opening to the lower opening of the funnel, an encapsulation (4) enclosing the funnel from the lower edge of the funnel to the upper edge of the funnel, such that a ventilation space (5) is defined within the one-piece body between the encapsulation and the funnel.

Abstract: A constellation of Earth-orbiting spacecraft includes a first spacecraft disposed in a first orbit and a second spacecraft disposed in a second orbit. Each of the first orbit and the second orbit is substantially circular with a radius of approximately 42,164 km. The first orbit and the second orbit have a respective inclination with respect to the equator within a range of 5° to 20°. The first orbit has a first right ascension of ascending node (RAAN1) and the second orbit has a second RAAN (RAAN2) of approximately RAAN1+90°.

Abstract: Techniques for performing spacecraft rendezvous and/or docking include operating a first orbiting spacecraft, the first spacecraft including a sensor arrangement, a first processor and a first inter-satellite link (ISL) arrangement and performing one or both of a rendezvous operation and a docking operation with the first spacecraft and a second orbiting spacecraft, the second spacecraft including one or more actuators. The performing one or both of the rendezvous operation and the docking operation includes determining a pose and pose rate of the second spacecraft relative to the first spacecraft using observations made by the sensor arrangement and determining a desired approach trajectory for the second spacecraft.

Abstract: A low-earth orbit (LEO) satellite includes a global positioning receiver configured to receive first signaling from a first plurality of non-LEO navigation satellites. An inter-satellite transceiver is configured to send and receive inter-satellite communications with other LEO navigation satellites. At least one processor is configured to execute operational instructions that cause the at least one processor to perform operations that include: determining an orbital position of the LEO satellite based on the first signaling; and generating a navigation message based on the orbital position. A navigation signal transmitter configured to broadcast the navigation message to at least one client device, the navigation message facilitating the at least one client device to determine an enhanced position of the at least one client device based on the navigation message and further based on second signaling received from a second plurality of non-LEO navigation satellites.

Abstract: An apparatus for launching, recovering, transporting, and storing vehicles is disclosed. The apparatus stabilizes the vehicle while it is in operation or inactive and has a frame connected to at least one stabilizer. In certain configurations, the apparatus stabilizes the vehicle without using the vehicle's onboard landing gear. The apparatus may also include at least one pad connected to the apparatus.

Abstract: A spacecraft includes a structural interface adapter for mating to a launch vehicle, an aft surface disposed proximate thereto, and a forward surface disposed opposite thereto, a main body structure, including a plurality of sidewalls, disposed between the aft surface and the forward surface and a plurality of unfurlable antenna reflectors. At least one unfurlable antenna reflector is configured to be illuminated, in an on-orbit configuration, by a respective feed array. The spacecraft is reconfigurable from a launch configuration to the on-orbit configuration. In the launch configuration, the at least one unfurlable antenna reflector is disposed, undeployed, forward of the respective feed array, proximate to and outboard of one of the plurality of sidewalls. In the on-orbit configuration, the forward surface is substantially earth-facing and the at least one unfurlable antenna reflector is disposed, deployed, so as to be earth-facing from a position aft of the respective feed array.

Abstract: A satellite system that includes a gateway, a satellite, and a user terminal. The gateway determines a modulation scheme based on a function of uplink and downlink signal quality and a defined relationship between the downlink modulation to the uplink modulation. The satellite includes an input demodulator configured to apply an input modulation and coding (modcod) scheme; an output modulator configured to apply an output modcod scheme; and an output modcod scheme selector configured to select an output modcod scheme for the output modulator based on the input modcod scheme according a predetermined relationship between input modcod schemes and output modcod schemes. The user terminal providing the gateway a measure of downlink signal quality.

Abstract: In some aspects, there is provided a method for determining a rotational orientation of an object in an image. The image depicts the object in a scene. The method includes providing images depicting the object in the scene to a trained statistical model. The images depict the scene of the image at different rotation angles. A rotation angle of a respective image corresponds to a potential rotational orientation of the object depicted in the respective image. The method further includes, in response to the providing, receiving, for each image, a confidence score indicating a likelihood generated by the trained statistical model that the object is at the potential rotational orientation corresponding to the rotation angle of the respective image. The method further includes determining the rotational orientation of the object in the image based at least in part on an analysis of the confidence scores and respective potential rotational orientations.

Abstract: A spacecraft includes a body, a plurality of separate units, and a first auxiliary radiator panel. The body includes a plurality of sidewalls, at least a first sidewall of the plurality of sidewalls including an outboard-facing radiator surface having optical solar reflectors disposed thereon. A first subset of the plurality of units is thermally coupled with the outboard-facing radiator surface of the first sidewall. A second subset of the plurality of units is thermally coupled with the first auxiliary radiator panel and is isolated from at least conductive thermal heat transfer with the outboard-facing radiator surface of the first sidewall. The first subset of units is spatially proximate to the second subset of units and is configured to operate in a first temperature range. The second subset of units is configured to operate in a second temperature range, the second temperature range being different from the first temperature range.

Abstract: Technique for altering a client spacecraft's rotational rate including the precise positioning of a servicing spacecraft in close proximity of a client spacecraft, alignment of a fluid release output device on the servicing spacecraft that imparts a force on the client spacecraft by means of fluid release, and subsequent use of the fluid release output device to mitigate tumbling of the client spacecraft. This allows the servicing spacecraft to slow the rotation of a tumbling client spacecraft in order to perform additional servicing operations.

Abstract: Systems and methods for satellite Synthetic Aperture Radar (SAR) artifact suppression for enhanced three-dimensional feature extraction, change detection, and/or visualizations are described. In some aspects, the described systems and methods include a method for suppressing artifacts from complex SAR data associated with a scene. In some aspects, the described systems and methods include a method for creating a photo-realistic 3D model of a scene based on complex SAR data associated with a scene. In some aspects, the described systems and methods include a method for identifying three-dimensional (3D) features and changes in SAR imagery.