Abstract: Spacecraft propulsion systems and methods featuring a first storage tank containing a metallic propellant and a second storage tank containing a non-metallic propellant are provided. A selected one of the metallic propellant and the non-metallic propellant is supplied to an electric propulsion thruster, depending on an operational mode of the spacecraft. The metallic propellant is stored at a relatively high density, while the non-metallic propellant is stored at a lower density than the metallic propellant. Moreover, the non-metallic propellant is preferably utilized to produce thrust through the electric propulsion thruster during operational maneuvers, while the metallic propellant is reserved for producing thrust through the electric propulsion thruster during end-of-life, such as deorbiting, maneuvers.

Abstract: A communication system has a first satellite, a second satellite, and a physical or wireless connection connecting the first satellite with the second satellite. The satellites each communicate with a different ground station and ensure continuous communication during handover.

Abstract: A method, apparatus, and system for controlling a device. A device control system comprises a computer system and a controller in the computer system. The controller receives internal sensor data for a group of internal parameters generated by an internal sensor system that senses the group of internal parameters within the device that relate to an operation of the device and receive external sensor data for a group of external parameters generated by an external sensor system that senses the group of external parameters in an environment around the device that relate to the operation of the device. The controller sends the internal sensor data and the external sensor data for an analysis with aggregated internal sensor data and aggregated external sensor data for devices of a same class as the device to generate results. The controller controls the operation of the device based on the results of the analysis.

Abstract: The subject of the invention is a detonation rocket engine comprising an annular detonation chamber (5) connected to the Aerospike nozzle (4) and lines (2, 3) for supplying propellant components connected to the detonation chamber (5). The detonation chamber (5) has a bottom (9) connecting the inner wall (10) and the outer wall (11) between which the outlet (6) is formed. At the outlet (6) of the detonation chamber (5) there are at least three evenly distributed centring elements (1) connecting the inner wall (10) and the outer wall (11) of the detonation chamber (5), with cooling channels (7) connected to one of the lines (2, 3) supplying the propellant components to the detonation chamber (5).

Abstract: An actively controlled deployment mechanism with a plurality of centrally controlled rollers as well as outer rollers that are configured to apply a pressure across the compactable portion of the deployable structure. The active control can be implemented though control motors and control loops programed to control the rotation of the rollers. Additionally, the actively controlled deployment reduces the need for bulky and heavy booms with high tension guide wires for the deployment of the structure.

Abstract: Systems and methods for identifying and removing a tracking capability from an external domain that performs a tracking activity on a host web page. Tracking capabilities of an external domain may be removed by altering web requests and/or responses to API calls. Once these tracking capabilities of the external domain have been removed, the altered web requests and/or altered responses to API calls may be transmitted to a web browser and/or entity making the API call thereby protecting user privacy while allowing the external domain to interact with the host web page.

Abstract: An intelligent control gas suction-type electric propulsion system applicable to multi-flow regimes: an ultra-low orbit rare gas is used as a working medium for attitude orbit control and resistance compensation propulsion, the gas is collected and inputted into an intelligent feedback pressurization system by means of a parabolic gas intake duct, intelligent feedback and pressurization are performed on the gas working medium by a molecular pump and a gas pump and the medium is stored in a working fluid storage tank so as to supply a hybrid thruster system that consists of seven sets of electric thrusters to generate thrust, which may achieve multiple thrust modes, and achieve the purpose of attitude orbit control and resistance compensation.

Abstract: A spacecraft refueling and storage system comprising a first tank and a second tank for storing propellant, a rotatable shaft to which the first and second tanks are mounted for rotating the first and second tanks about an axis of the shaft, and a drive motor for rotating the shaft so that upon rotation of the first and second tanks, liquid propellant is separated from gas in the propellant and settled to an outer portion of the first and second tanks.

Abstract: An example system for extraterrestrial deployment of a flexible membrane surface includes a flexible membrane having a periphery and an interior. The flexible membrane is rolled about a roll axis into a cylindrical geometric shape in an undeployed state. A payload base has extendable radial booms, wherein the distal end of each extendable radial boom is attached to the periphery of the flexible membrane and the interior of the flexible membrane is free of attachment to the extendable radial booms. The payload base and the extendable radial booms are positioned to one side of the flexible membrane along the roll axis. The extendable radial booms are configured to extend orthogonally to the roll axis from the payload base to unroll the flexible membrane about the roll axis to form the flexible membrane surface in a deployed state, wherein the roll axis is substantially orthogonal to the flexible membrane surface.

Abstract: A vapor-pressure driven micro pump system comprising an enclosure of a supporting structure; a first chamber having a first volatile material as a propellant with a plurality of exit nozzles; a second chamber have a second volatile material inside of a collapsible diaphragm which separates the first and the second chambers within said enclosure, wherein a vacuum at the plurality of exit nozzles causes vaporization of said propellant, which is compensated and displaced by vapor of said second volatile material at a substantially constant pressure by moving of said collapsible diaphragm. The vapor-pressure driven pump system is useful for various situations, especially in a gravity-free environment in space exploration.

Abstract: The invention is methods for landing rockets with precision using deep reinforcement learning for control. Embodiments of the invention are comprised of three steps. First, sensors collect data about the rocket's physical landing environment, passing information to rocket's database and processors. Second, the processors manipulate the information with a deep reinforcement learning program to produce instructions. Third, the instructions command the rocket's control system for optimal performance during landing.

Abstract: A launch vehicle to transport at least one payload into an earth orbit, wherein the launch vehicle comprises a plurality of solar cells on its outer surface. Furthermore, a manufacturing method for a launch vehicle with a plurality of solar cells on its outer surface and a transport method for at least one payload using a launch vehicle with a plurality of solar cells on its outer surface are provided.

Abstract: The disclosed propulsion system of a space vehicle and the methods of operating the propulsion system use a microwave energy source to heat propellant in a propellant chamber that pierces and traverses a waveguide carrying the microwave energy. In some implementations, the microwave energy ionizes and further heats the propellant in the propellant chamber. The partially ionized and heated propellant may exit the propellant chamber via a nozzle to generate thrust.

Abstract: Microwave ignition systems with a launcher affixed to or located within a gun spindle. Use of a planar, impedance matched system that include a launcher affixed to or located within a gun spindle and a receiver affixed the propelling charge, where the receiver is engineered and impedance matched to efficiently deposit energy into the receiving igniter material, may result in significantly better timing and reliability than conventional mechanical gun ignition systems and other microwave ignition systems. A pressure-tolerant feed through system can route the microwave energy to the inside of the breech.

Abstract: A cooling system and cooling method for a fuel cell onboard a vehicle, wherein the fuel cell includes a gasifier configured to expand liquid hydrogen to gaseous hydrogen for feed to the fuel cell, the cooling system including a coolant system sized for less than peak power operation of the vehicle; and an auxiliary coolant system configured to provide supplemental cooling to the fuel cell, wherein the supplemental coolant system is configured to by-pass coolant to the gasifier and employ heat of gasification of the liquid hydrogen to provide supplemental cooling for the fuel cell during peak vehicle operation.

Abstract: A multi-mode thruster system for use in a spacecraft includes a microwave source; a cavity coupled to the microwave source and including a first inlet to receive a first fluid and a second inlet to receive a second fluid; and a nozzle provided at one end of the cavity. The thruster operates in a microwave electrothermal thruster (MET) mode to (i) generate a standing wave in the cavity using the microwave source and (ii) raise a temperature of the first fluid to generate a first hot gas that exits the cavity via the nozzle to generate thrust. The thruster operates in a chemical propulsion mode to (i) produce a reduction-oxidation reaction between the first fluid and the second fluid and (ii) generate a second hot gas that exits the cavity via the nozzle to generate thrust.

Abstract: A device for securing at least one thermal protection mat to an internal fixed structure of a nacelle includes at least one removable fastening element for securing the at least one thermal protection mat to the internal fixed structure designed to be arranged between the mat and the internal fixed structure. The device includes at least one gripping element connected to the removable fastening element to project relative to the mat on the opposite side to the internal fixed structure, the removable fastening element being intended to be detached from the internal fixed structure by a pulling force exerted on the gripping element.

Abstract: Spacecraft servicing devices or pods and related methods may be configured to be deployed from a carrier spacecraft and include at least one spacecraft servicing component configured to perform at least one servicing operation on the target spacecraft. The spacecraft servicing devices may be configured to be transported from an initial orbit to another orbit after the spacecraft servicing device is deployed from the carrier spacecraft.

Abstract: Methods and systems for mitigating or reducing the risk of an electrostatic discharge due to static charge differentials between a first spacecraft and a second spacecraft as the first spacecraft approaches the second spacecraft may be accomplished using a passive electrostatic discharge mitigation device. In some embodiments, mitigation of static potential between the first spacecraft and the second spacecraft may be actively accomplished by an electric propulsion system provided on the first spacecraft. In some embodiments, mitigation may be provided by both actively and passively mitigating static potential between the first spacecraft and the second spacecraft.

Abstract: The invention relates to a method for transferring a space payload from a first orbit to a second orbit. Said method is characterized in that the space payload, moving about the first orbit, is attached to a removable orbital towing assistance device including at least one fuel pouch. Said method includes the steps of: —attaching (E1) an orbital transfer vehicle to the removable orbital towing assistance device; and —transferring (E2) the space payload and the removable orbital towing assistance device to the second orbit by means of the orbital transfer vehicle. The invention also relates to a removable orbital towing assistance device intended for a space payload and an orbital transfer vehicle and enabling direct supply of fuel to the vehicle and/or the space payload.

Abstract: Methods and systems for mitigating or reducing the risk of an electrostatic discharge due to static charge differentials between a first spacecraft and a second spacecraft as the first spacecraft approaches the second spacecraft may be accomplished using a passive electrostatic discharge mitigation device. In some embodiments, mitigation of static potential between the first spacecraft and the second spacecraft may be actively accomplished by an electric propulsion system provided on the first spacecraft. In some embodiments, mitigation may be provided by both actively and passively mitigating static potential between the first spacecraft and the second spacecraft.

Abstract: The present application relates to optical-mechanical systems and methods for moving a solid object by applying conservation of angular momentum to a configuration of a laser light beam that emanates from the solid object. The system includes a rotatable housing and an axially movable laser light source coupled to the housing and configured to emit a first light beam along a first path. The system can include a first beam splitter disposed along the first path for splitting the first light beam into a second light beam and a third light beam. The system can cause the third light beam to travel in a closed path, as an approximation of a circular path of initial radius, and of decreasing radius. The system can further include a second beam splitter, axially movable first, second and third mirrors, and a third beam splitter disposed at one end of the housing.

Abstract: A field of feeding propellant gas to ion thrusters, and in particular a method of feeding an ion thruster with propellant gas coming from a pressurized tank via a feed circuit including an on/off valve and, in succession downstream from the on/off valve, a high pressure restrictor, a buffer tank, and at least one low pressure restrictor. The method includes the steps of calculating a pressure setpoint for the buffer tank as a function of a flow rate setpoint, calculating the difference between the pressure setpoint for the buffer tank and a pressure measured in the buffer tank, calculating a setpoint for the opening time of the on/off valve as a function of the difference and of a pressure inside the pressurized tank, and opening the on/off valve in compliance with the opening time setpoint.

Abstract: A method for actuating an electromagnetic valve in a fluid system includes, for a specified first time period, a switching current with a specified first amplitude is applied, the switching current switching the electromagnetic valve from a rest state into a switching state. After the specified first time period expires, a holding current with a specified second amplitude is applied, the holding current holding the electromagnetic valve in the switching state. The first amplitude of the first switching current is greater than the second amplitude of the holding current.

Abstract: A hypersonic aircraft having a homopolar motor with high temperature superconducting (HTS) non-insulated (NI) coil magnets is described. In some implementations, the HTS NI coil magnets can have a graded resistance design. In some implementations, the HTS NI coil magnets can include a series of stacked coils, each of the series of coils comprising multiple turns having turn-to-turn resistance, where the turn-to-turn resistance of the series of coils is graded coil-to-coil across the magnet. In some implementations, the HTS NI coil magnets can include an NI coil comprising multiple turns and two or more thermal barriers each disposed between two adjacent turns of the coil, where an electrically conductive portion of one of the thermal barriers does not overlap with an electrically conductive portion of a different adjacent one of the thermal barriers. Some implementations can include a disk-type homopolar motor/generator including one or more HTS NI coil magnets.

Abstract: Systems and methods that optimize landing performance are provided. The system determines a number, N, of equipment configurations (a combination of a brake setting and a thrust reverser configuration) supported by the aircraft. The system determines a deceleration airspeed to achieve a target taxi speed and, for each of the N equipment configurations, determines a respective deceleration distance. In various embodiments, the system further updates the deceleration distances by one or more of a brake's condition, aircraft historical data, brake warranty and life cycle data, and environmental conditions. The deceleration distances are used to identify a number P of exit-ways that can be used at the runway. Total costs (including brake usage and fuel cost) for each of the P exit-ways is determined, and the equipment configuration that delivers the lowest total cost delivers the optimize landing performance.

Abstract: Devices, methods and systems for minimizing the probability of a collision between an aircraft and a floating platform are described. The device may include a processor in communication with a memory. The processor is configured to obtain a flight-path vector of an aircraft; determine a probability related to a plurality of flight-paths of a floating platform over a period of time based on operating parameters for the floating platform and weather data; and determine, based on the flight-path vector and the probability related to the plurality of flight-paths of the floating platform, a time and/or a location for launch or recovery of the floating platform that minimizes a probability of a collision between the aircraft and the floating platform while the floating platform is in flight.

Abstract: A system and method for assisted EVA self-return is provided herein. The system estimates a crewmember's navigation state relative to a fixed location, for example on an accompanying orbiting spacecraft, and computes a guidance trajectory for returning the crewmember to that fixed location. The system may account for safety and clearance requirements while computing the guidance trajectory. According to at least one embodiment, the system actuates the crewmember's safety jetpack to follow the prescribed trajectory to the fixed location. In another embodiment, the system provides the crewmember with a directional cue (e.g., a visual, auditory, or tactile cue) corresponding to the prescribed trajectory back to the fixed location. The system may be activated by the crewmember or remotely by another crewmember and/or system.

Abstract: Methods of laser powering unmanned aerial vehicles (UAV) with heat engines are disclosed. The laser powered heat engines are used in conjunction with devices for absorbing laser optical radiation, turning the laser optical radiation into heat, supplying the heat to a working fluid of the heat engine and harvesting mechanical work from expanding working fluid in the heat engine.

Abstract: A docking node transporter tug is disclosed. The tug can dock with various spacecraft to allow access by people between the spacecraft. Further, the tug may dock with other specialty tugs to form a custom system.

Abstract: A spacecraft capture tug is disclosed. The tug can capture objects in space, including spacecraft and satellites, using an engagement system and provide propulsion for the object to a point of servicing, repair, or other desired operations. Further, the tug may dock with other specialty tugs to form a custom transport system.

Abstract: An orbit attitude control device includes a plurality of nozzles for injecting combustion gas supplied from a combustion chamber, and a control section configured to calculate nozzle opening degree correction values so that a deviation between a detection value of the pressure of the combustion chamber and a command value becomes smaller. The control section is configured to calculate a total correction value so that the deviation between the detection value and the command values becomes smaller. A total value T1 for first group nozzles and a total value T2 for second group nozzles are calculated. The total correction value is distributed to the opening degree correction values for the first group nozzles with a ratio of T2/(T1+T2) and to the opening degree correction values for the second group nozzles with a ratio of T1/(T1+T2).

Abstract: An orbit attitude control device includes a plurality of nozzles for injecting combustion gas supplied from a combustion chamber, and a control section configured to calculate nozzle opening degree correction values so that a deviation between a detection value of the pressure of the combustion chamber and a command value becomes smaller. The control section is configured to calculate a total correction value so that the deviation between the detection value and the command values becomes smaller. A total value T1 for first group nozzles and a total value T2 for second group nozzles are calculated. The total correction value is distributed to the opening degree correction values for the first group nozzles with a ratio of T2/(T1+T2) and to the opening degree correction values for the second group nozzles with a ratio of T1/(T1+T2).

Abstract: A standard transit tug is disclosed. The tug can dock with spacecraft to provide propulsion for the spacecraft. Further, the tug may dock with other specialty tugs to form a custom transport system.

Abstract: A solar generator tug is disclosed. The tug can dock with spacecraft to provide power for the spacecraft. Further, the tug may dock with other specialty tugs to form a custom transport system.

Abstract: The application describes a wind turbine having a control method and controller for performing predictive control of a wind turbine generator. Based on the measured instantaneous wind speed, it is known to provide control signals to a wind turbine in order to control the pitch of the wind turbine rotor blades and the speed of the generator. However, it is difficult using instantaneous wind speed measurements to achieve smooth control, due to finite response speeds of the associated electro-mechanical systems, as well as the constantly changing control system inputs. The predictive control system described in the application assumes a model of generator speed based on the values of the incident wind speed v(t) and the values of a control signal u(t) output to the wind turbine in a feed forward loop. Here, the control signal can be for one or more of controlling either the power setting of the generator, or the pitch angle of the rotor blades.

Abstract: A space debris remover aiming to remove a space debris object in earth orbits. Angular thrust calculation unit calculates angular thrust. Radial thrust calculation unit calculates radial thrust based on the angular thrust, estimated angular momentum and estimated space debris mass. A foam bonding mechanism connects the space debris remover and the space debris object. A space debris removal controller calculates firing time, and sends a space debris removal control signal comprising the radial thrust, the angular thrust and the firing time. A plurality of first stage thrusters generate the radial thrust and the angular thrust after the firing time. After the stage separator separates a first stage and a second stage of the space debris remover, a plurality of second stage thrusters generate the radial thrust and the angular thrust, and propel the space debris object towards the sun.

Abstract: Systems and methods for launching space vehicles into orbit involve placing a space vehicle into a protective container. The protective container is placed on a sled in a maglev tunnel and then launched into the atmosphere. Once the protective container reaches a certain height the space vehicle is released from the protective container and the space vehicle then proceeds to the desired orbital position.

Abstract: A space debris remover aiming to remove a space debris object in earth orbits. Angular thrust calculation unit calculates angular thrust. Radial thrust calculation unit calculates radial thrust based on the angular thrust, estimated angular momentum and estimated space debris mass. A foam bonding mechanism connects the space debris remover and the space debris object. A space debris removal controller calculates firing time, and sends a space debris removal control signal comprising the radial thrust, the angular thrust and the firing time. A plurality of first stage thrusters generate the radial thrust and the angular thrust after the firing time. After the stage separator separates a first stage and a second stage of the space debris remover, a plurality of second stage thrusters generate the radial thrust and the angular thrust, and propel the space debris object towards the sun.

Abstract: A spacecraft system and method includes a platform with a dock and an umbilical payout device. A robot is connected to an umbilical paid out by the umbilical payout device and is repeatedly deployable from the dock. The robot includes one or more imagers, an inertial measurement unit, and a plurality of thrusters. A command module receives image data from the one or more robot imagers and orientation data from the inertial measurement unit. An object recognition module is configured to recognize one or more objects from the received image data. The command module determines the robot's orientation with respect to an object and issues thruster control commands to control movement of the robot based on the robot's orientation. The combination of the space platform and robot on umbilical line can be used for towing another object to different orbital location, inspection including self-inspection of the robot carrying platform and for robotic servicing.

Abstract: A system for enhancing movement of a multi-engine flight vehicle about either of pitch or yaw axes. The engines are oriented relative the centerline of the flight vehicle and a portion of the engines have a positive tangential cant and the remaining engines have a negative tangential cant. The attitude of the flight vehicle about the pitch, yaw and roll axes is controlled by differentially throttling the multi-engine flight vehicle.

Abstract: A system and methods are provided for combining systems of an upper stage space launch vehicle for enhancing the operation of the space vehicle. Hydrogen and oxygen already on board as propellant for the upper stage rockets is also used for other upper stage functions to include propellant tank pressurization, attitude control, vehicle settling, and electrical requirements. Specifically, gases from the propellant tanks, instead of being dumped overboard, are used as fuel and oxidizer to power an internal combustion engine that produces mechanical power for driving other elements including a starter/generator for generation of electrical current, mechanical power for fluid pumps, and other uses. The exhaust gas from the internal combustion engine is also used directly in one or more vehicle settling thrusters. Accumulators which store the waste ullage gases are pressurized and provide pressurization control for the propellant tanks.

Abstract: A microsatellite comprising a propulsion module for moving and/or pointing the microsatellite and an imaging device (100) mounted to the propulsion module. Furthermore there is a fuel supply (210) located within the imaging device (100).

Abstract: A Kinetic Energy Storage and Transfer (KEST) vehicle and target vehicle kinetic energy transfer method are provided. The KEST vehicle is configured to transfer kinetic energy to the target vehicle, propelling the target vehicle into a higher orbit or beyond the Earth. This is accomplished by a catching mechanism that contacts the target vehicle. The catching mechanism may also include a braking mechanism configured to accelerate the target vehicle, and thus slow the KEST vehicle, as the catching mechanism and target vehicle travel along one or more tethers of the KEST vehicle. Alternatively, the catching mechanism may be attached to an end of the one or more tethers and be configured to slow the target vehicle as the one or more tethers bend.

Abstract: The invention relates to a surface section that is exposed to an ion flow, in particular for a drive arrangement in a spacecraft, comprising an electrostatic ion accelerator arrangement. According to the invention, in order to reduce erosion, an intermediate potential energy surface is provided, the surface being advantageous in that it allows a magnetic field that is essentially parallel to the surface section to be formed.

Abstract: A space vehicle (1), in particular a satellite or a probe, capable of executing high-thrust maneuvers or high specific impulse maneuvers. According to the invention, the space vehicle (1) includes at least one solid-propellant chemical thruster (10-14) and at least one electric thruster (20-24).

Abstract: A standard transit tug is disclosed. The tug can dock with spacecraft to provide propulsion for the spacecraft. Further, the tug may dock with other specialty tugs to form a custom transport system.

Abstract: A multiple space vehicle launch system that may be adapted to be disposed within a payload region of a launch vehicle fairing is disclosed. The vehicle launch system may include a first space vehicle including a first core structure and a second space vehicle including a second core structure. The second core structure may be oriented relative to the first space vehicle such that when placed within a fairing, a launch load is transmitted from the first core structure of the first space vehicle and is borne by the second core structure of the second space vehicle. The first space vehicle and the second space vehicle each include respective propulsion units.

Abstract: A thruster includes a reactor including bundled metal elongate members therein configured to disassociate an introduced propellant flowing between the members, a combustor for combusting products of the disassociation, and a nozzle downstream of the combustor.

Abstract: This invention relates to a propulsion bay to be transported, at least temporarily, in a space launch vehicle and comprising an adapter that co-operates with at least one system located, at least temporarily, on board the bay, said system comprising an electrical power supply. The bay is characterized in that it also comprises at least one electric space propulsion engine that can be powered by the power supply of the system.