1717results about "Cosmonautic component separation" patented technology

Systems and methods are provided for swapping the battery on an unmanned aerial vehicle (UAV). The UAV may be able to identify and land on an energy provision station autonomously. The UAV may take off and / or land on the energy provision station. The UAV may communicate with the energy provision station. The energy provision station may store and charge batteries for use on a UAV.

Compression exerted by a clamp collar integral to a molded plastic flexible shaft coupling accommodating angular shaft misalignment upon the circumferential outward surface of a radially compressible molded sleeve insert located in a shaft bore of the flexible coupling facilitates engagement of a reduced diameter shaft located within the sleeve insert. Use of a sleeve insert having a different internal diameter similarly accommodates flexible coupling to another diameter shaft. Utilization of sleeve inserts selected from a plurality of different reduced internal diameters and exterior diameters sized to fit either shaft bore enables flexible coupling of both multiple matched shaft diameters and an exponential number of differing opposed shaft diameter combinations within a range limited by the flexible coupling shaft bores. Radial compressibility of the sleeve insert is facilitated by the utilization of resilient material, interruption of the circumferential extension by a longitudinal slot, or both. Utilization of congruent longitudinally uniform, radially non-uniform, flexible coupling shaft bore and insert sleeve outward surfaces provides additional torque transmission capacity.

The present invention provides a docking mechanism that is capable of interfacing with Apogee Boost Motors (ABM) including Liquid Apogee Motors (LAM) of the satellite being captured to allow a servicing spacecraft to dock with a satellite that has no special docking features. The docking mechanism includes a compliant probe with spring-loaded fingers at its tip which is inserted into the LAM by the approach motion of the servicer spacecraft. When the probe tip has passed beyond the throat of the LAM the finger extend to trap the LAM on the probe thus achieving capture. The rigidization stage of docking is achieved by retracting the probe to pull the two spacecraft together. The docking mechanism includes abutment pads which interface with the launcher interface ring of the client spacecraft. With the preload applied by the probe retraction mechanism reacted through these pads, the docking mechanism can maintain a rigid docked interface under loads induced by thrusting maneuvers of the combined spacecraft or loads induced by on-orbit servicing operations.

A closed-loop LRF pointing technology to measure the range of a target satellite from a chaser satellite for rendezvous is provided that includes several component technologies: LOS angle measurements of the target satellite on a visible sensor focal plane and the angles' relationships with the relative position of the target in inertial or LVLH frame, a relative navigation Kalman filter, attitude determination of the visible sensor with gyros, star trackers and a Kalman filter, pointing and rate commands for tracking the target, and an attitude controller. An analytical, steady-state, three-axis, six-state Kalman filter is provided for attitude determination. The system and its component technologies provide improved functionality and precision for relative navigation, attitude determination, pointing, and tracking for rendezvous. Kalman filters are designed specifically for the architecture of the closed-loop system to allow for pointing the laser rangefinder to a target even if a visible sensor, a laser rangefinder, gyros and a star tracker are misaligned and the LOS angle measurements from the visible sensor are interrupted.

A system for capturing and docking an active craft to a passive craft has a first docking assembly on the active craft with a first contact member and a spike projecting outwardly, a second docking assembly on the passive craft having a second contact member and a flexible net deployed over a target area with an open mesh for capturing the end of the spike of the active craft, and a motorized net drive for reeling in the net and active craft to mate with the passive craft's docking assembly. The spike has extendable tabs to allow it to become engaged with the net. The net's center is coupled to a net spool for reeling in. An alignment funnel has inclined walls to guide the net and captured spike towards the net spool. The passive craft's docking assembly includes circumferentially spaced preload wedges which are driven to lock the wedges against the contact member of the active craft. The active craft's docking assembly includes a rotary table and drive for rotating it to a predetermined angular alignment position, and mating connectors are then engaged with each other. The system may be used for docking spacecraft in zero or low-gravity environments, as well as for docking underwater vehicles, docking of ancillary craft to a mother craft in subsonic flight, in-flight refueling systems, etc.

A recovery craft with a coupling arrangement captures, engages, and transports a defective or expended non-maneuverable spacecraft. The coupling arrangement includes a coupling mast (e.g. a telescoping pipe), a releasable rigid mount that secures one end of the coupling mast to the recovery craft, and spreader arms that are radially spreadable from the other end of the coupling mast. The spreader arms are initially radially inwardly retracted, and are inserted into an interface ring of the spacecraft. Then the spreader arms are radially outwardly extended to engage behind a protruding rim of the interface ring. A spring braces against the spacecraft. Next, the rigid mount is released and the coupling mast remains connected to the recovery craft only by a tension-transmitting cable. Thrusters of the recovery craft are activated to tow the spacecraft to a new target position, orbit, or trajectory.

The invention relates to a large-tolerance docking acquisition device, in particular to a large-tolerance docking acquisition device focused on a space large mechanical arm and rendezvous and docking. The large-tolerance docking acquisition device is provided for solving the problems that the existing cone-pole-type acquisition mechanism cannot acquire a small-inertia unpowered target freely floating in the space, and the allogeneic isomorphic peripheral docking mechanism has a small tolerance range. The large-tolerance docking acquisition device comprises an acquisition mechanism and an interfacing mechanism; the acquisition mechanism is arranged on a wrist joint of an active satellite or a space mechanical arm, and the interfacing mechanism is mounted on an acquired satellite or acquired load. The large-tolerance docking acquisition device is used for acquiring the small-inertia unpowered target freely floating in the space or performing the allogeneic isomorphic peripheral docking.

This invention solves problems associated with prior-art soft-dock mechanisms by placing all active components of a soft-dock system on the chaser side of the mechanism, leaving the target side of the mechanism completely passive (i.e., requiring no power expenditure or self-actuated moving parts to operate). In particular, the active components are supported on the end of a flexible cable attached to the probe, or chaser, side of the device. These components act as a sort of spring-loaded “trap.” Once the end of the probe passes into a receptacle on the target side, the mechanism is triggered, engaging it in such a way that it can no longer be pulled out of the receptacle until it is reset. The soft-docking cable may be replaced with a rigid, semi-rigid or jointed post that is used to bring a capture mechanism into engagement with its corresponding receptacle or receiving structure. The magnetic end effector may also be implemented as an electro-magnet, which requires power to maintain the holding force, or a permanent magnet, which captures a target without power. The main target cone may be either a metallic cone: or a non-metallic cone constructed of fabric, plastic, or other flexible material.

An ordnance firing system is disclosed that includes a reusable electronics module and an ordnance module, each enclosed in a separate, sealed housing. The electronics module housing encloses firing electronics for electrically triggering initiation of a detonator in the ordnance module. The electronics module detachably connects to the ordnance modules via a connector which extends away from the electronics module housing. The housing of the ordnance module is constructed to be blast-resistant to prevent detonation of the detonator from rendering the electronics module inoperable.

An iso-truss structure (10) includes at least two helical components (30, 32) and at least one reverse helical component (34) attached thereto with opposing angular orientations. Each helical and reverse helical component preferably includes at least four elongate, straight segments (22) rigidly connected end to end in a helical configuration forming a single, substantially complete rotation about a common axis (14) forming a first square cross section. The structure may further include at least two rotated helical components (80, 92) and at least one rotated reverse helical component (98) which are rotated with respect to the helical and reverse helical components forming a second square cross section, rotated with respect to the first. The structure may be straight, curved, flexible, or form angles.

A system and method for refueling unmanned aerial vehicles. The system is adapted to refuel a first unmanned aerial vehicle from a second unmanned aerial vehicle and includes an arrangement for flying the first and second vehicles to proximity within a predetermined range and for connecting an umbilical from the second vehicle to the first vehicle in flight. In the illustrative embodiment, the arrangement for connecting includes a targeting system for electromagnetically detecting a refueling receptacle on the first vehicle. The targeting system includes a first coil around a refueling receptacle on the first vehicle. A seeker is disposed at a first end of said umbilical on the second vehicle. The seeker includes three detector coils adapted to detect a magnetic signal from the first coil around the receptacle on the first vehicle. The coils are mounted such that the detector coils point in different directions. The outputs of the coils are processed to determine the direction and range to the UAV from the tanker UAV.

A canisterized satellite dispenser includes one or more of: a pair of guide rails that eliminate the requirement of a rectangular profile for the satellite; a preload system that secures the canisterized satellite during transport and launch, and releases to deploy the canisterized satellite; a constant-force spring to provide a uniform and predictable dispensing force; an external rectangular profile in each dimension; and internal support surfaces that simplify the design of canisterized satellites, particularly those with deployable components. Each canisterized satellite includes a pair of opposing flanges on a lower portion of the satellite that ride in a channel formed by the dispenser's guide rails and restraining flanges; no other support constraints are imposed. During travel and launch, the satellite flanges are held against the restraining flanges, rigidly fixing the satellite to the dispenser until the satellite is deployed.

The present invention provides a method, system and apparatus for robotic refueling of satellites. The system may include a dedicated refueling satellite launched directly from either earth, or alternatively it could be launched from another larger mother spacecraft or space station in which the refueling satellite is ferried into space in the case of the larger space craft or it may be stored on the space station until needed from which it can be launched. The system includes a robotic arm, suitable tools which can be affixed to the end effector of the robotic arm required for accessing, opening and closing the fuel fill valve on the satellite being serviced, storage and retrieval stations on a tool caddy on which the tools and various fuel fill valve caps are stored. The system is under teleoperation by a remotely located operator, for example located on earth, in the mother station or in the space station. Cameras are included focussed on the robotic arm and end effector and images are transmitted to the operator to allow the operator to direct and control the refueling procedure. The system may also be configured to be operated autonomously under computer control.

The invention discloses a small-size spacecraft butt-joint mechanism, relates to a spacecraft butt-joint mechanism and aims at solving the problems that an existing small-size spacecraft butt-joint mechanism is fierce in collision, low in positioning precision, and large in control difficulty in the butt-joint process and unstable enough after butt-joint. A connecting sleeve and three arc-shaped enclosing plates are arranged between an upper driven disc and a lower driven disc; a first guide plate is arranged at one end of each arc-shaped enclosing plate; a second guide plate is arranged at the other end of each arc-shaped enclosing plate; the adjacent first guide plates and three second guide plates form a V-shaped positioning slot; three brackets are uniformly distributed between the upper drive disc and the lower drive disc along the same circumference; a stepped motor is fixed in the center of the lower drive disc; the stepped motor is connected with a lead screw; a threaded lifting disc is screwed on the lead screw; the lower end of each locking claw penetrates through a slide path and is hinged with the threaded lifting disc, and each locking claw is contacted with the surface of a roller; the upper end of each locking claw is arranged in the corresponding V-shaped positioning slot. The small-size spacecraft butt-joint mechanism disclosed by the invention is mainly used for capturing and butt-joint of a small-size spacecraft.

A method and apparatus for conducting proximity operations is disclosed. The method called inverser proximity operations includes maneuvering an active vehicle into general proximity to a target vehicle, transmitting from the target vehicle to the active vehicle data representing relative position and velocity between the target vehicle and the active vehicle, and maneuvering the active vehicle in accordance with the data to effect a desired proximity operation. Another method called distributed proximity operations includes maneuvering a carrier vehicle into general proximity to a target vehicle, releasing one or more active vehicles from the carrier vehicle, transmitting from the carrier vehicle to the active vehicle(s) data representing relative position and velocity between the target vehicle and the active vehicle, and maneuvering the active vehicle(s) in accordance with the data to effect a desired proximity operation. The proximity operations are similarly suitable for aeronautical systems, such as when using an active aircraft to re-fuel a target aircraft, such as a drone. Whether for space, air, or other environments, the proximity operations described can be used for re-fueling, repairing, and replacing components and / or systems.

An apparatus for launching and dispensing multiple satellites into orbit includes a central structure and a plurality of satellites releasably bundled around the central structure to form a cohesive, single payload mounted on an upper stage of a launcher rocket. The central structure includes a central tube and an adapter shell that fits onto the upper stage of the rocket. Each satellite has the shape of a partial cylinder segment of which an outer wall is a cylindrical segment shell. The satellites are bundled and secured together around the central structure by tension bands, whereby the cylindrical segment shells form a load-bearing complete outer cylindrical shell that surrounds the central tube. This structure provides sufficient stiffness and strength to withstand the loads that arise during launch, while having a reduced mass. The satellites themselves, rather than the central structure, bear the loads during launch. This allows the central structure to be of relatively lightweight construction.

An apparatus for orbital change of a satellite incorporates a capture mechanism for space debris and a tether connecting the capture mechanism to a satellite. The tether is extendable to position the capture mechanism relative to the space debris. A controller is employed for timed release of the space debris by the capture mechanism for orbital change by the satellite.

Apparatus and methods for performing satellite proximity operations such as inspection, recovery and life extension of a target satellite through operation of a “Satellite Inspection Recovery and Extension” (“SIRE”) spacecraft which can be operated in the following modes (teleoperated, automatic, and autonomous). The SIRE concept further consists of those methods and techniques used to perform certain (on-orbit) operations including, but not limited to, the inspection, servicing, recovery, and lifetime extension of satellites, spacecraft, space systems, space platforms, and other vehicles and objects in space, collectively defined as “target satellites”. The three basic types of SIRE proximity missions are defined as “Lifetime Extension”, “Recovery”, and “Utility”. A remote cockpit system is provided to permit human control of the SIRE spacecraft during proximity operations.

An In Orbit Transportation & Recovery System (IOSTAR™) (10) One preferred embodiment of the present invention comprises a space tug powered by a nuclear reactor (19). The IOSTAR™ includes a collapsible boom (11) connected at one end to a propellant tank (13) which stores fuel for an electric propulsion system (12). This end of the boom (11) is equipped with docking hardware (14) that is able to grasp and hold a satellite (15) and as a means to refill the tank (13). Radiator panels (16) mounted on the boom (11) dissipate heat from the reactor (19). A radiation shield (20) is situated next to the reactor (19) to protect the satellite payload (15) at the far end of the boom (11). The IOSTAR™ (10) will be capable of accomplishing rendezvous and docking maneuvers which will enable it to move spacecraft between a low Earth parking orbit and positions in higher orbits or to other locations in our Solar System.

An androgynous mating system for mating two exoatmospheric space modules comprising a first mating assembly capable of mating with a second mating assembly; a second mating assembly structurally identical to said first mating assembly, said first mating assembly comprising; a load ring; a plurality of load cell subassemblies; a plurality of actuators; a base ring; a tunnel; a closed loop control system; one or more electromagnets; and one or more striker plates, wherein said one or more electomagnets on said second mating assembly are capable of mating with said one or more striker plates on said first mating assembly, and wherein said one or more striker plates is comprised of a plate of predetermined shape and a 5-DOF mechanism capable of maintaining predetermined contact requirements during said mating of said one or more electromagnets and said one or more striker plates.

The invention discloses separable micro and nano-satellite configuration which comprises a satellite body, a plurality of separate release sub-satellites, mechanical net claws, a flexible solar cell array and a variable-structure mechanical arm. The separable micro and nano-satellite configuration has the advantages that the integral separable micro and nano-satellite configuration is in such a configuration form that a primary satellite carries the multiple sub-satellites, and functions of separately releasing and hovering the micro and nano-satellites can be realized; the efficiency of the solar cell array can be improved owing to combined configuration comprising flexible solar wings and the variable-structure mechanical arm; mechanical net claw mechanisms are carried on the micro and nano-satellites, so that a function of capturing space targets can be realized; the separable micro and nano-satellite configuration is provided with a plurality of stereoscopic vision imaging cameras, so that a stereoscopic monitoring function can be realized; attitude maneuver and attitude stabilization of the micro and nano-satellites can be implemented by the aid of a solid miniature propulsion technology, so that the sub-satellite separate release control ability can be improved; the separable micro and nano-satellite configuration is applicable to fifty-kilogram-level micro-miniature artificial earth satellites on near-earth orbits and is also applicable to micro-satellite networking, scientific detection aircrafts and relevant micro and nano-satellite space demonstration and experiments.

An automated docking system for space vehicles that includes a plurality of antennas on each of a target vehicle and a chase vehicle. A pseudo random code is transmitted via one of the antennas located on one of the vehicles and received by one of the antennas located on the other vehicle. The pseudo random code is then sent back to the original vehicle via transmissions from the plurality of antennas on the second vehicle and received by the plurality of antennas on the original vehicle. The distance from each of the antennas on the other vehicle to each of the antennas on the originating vehicle can be measured in this fashion. The antennas on each of the vehicles are located in a spaced-apart arrangement so that the angular orientation or attitude of each of the vehicles to each other can also be determined. A plurality of video cameras is provided on the exterior of one of the vehicles and video information from these cameras is transmitted to the other vehicle for display to operators in that vehicle. Commands between the vehicles can also be communicated.

Solar array assemblies and systems and methods for deploying solar cell arrays from a spacecraft. A solar cell panel assembly comprises a first flexible solar panel and a rotational member. A first extension assembly is disposed proximate to a first end of the rotational member and a second extension assembly is disposed proximate to a second end of the rotational member. The solar cell panel assembly further comprises a first support member and a second support member. The first end of the first support member is coupled to the first end of the rotational member and the first end of the second support member is coupled to the second end of the rotational member. The second ends of the support members are coupled to the first flexible solar panel. A tether assembly couples the extension assemblies to the support members.

The invention provides an under-actuated three-arm type non-cooperative target docking-capturing device, relating to a docking-capturing device. The invention solves the problem that: the prior non-cooperative target docking mechanism is easy to collide with satellite spray pipes in a docking process, thereby causing poor docking reliability. An under-actuated manipulator motor (18) is arranged on an upper end face of a lead screw nut plate (2) on a lifting mechanism; an output shaft of the under-actuated manipulator motor (18) is connected with a ball lead screw (23); the ball lead screw (23) is sleeved with a nut (21); a manipulator driving plate (17) is fixed on the nut (21); three manipulators are evenly arranged on a manipulator mounting platform (4); the other ends of three first coupling-connecting rods are respectively in rotary connection with the manipulator driving plate (17); and an umbrella-shaped buffer mechanism is positioned above and flexibly connected with the manipulator mounting platform (4). The docking-capturing device has the advantages of high docking reliability, wide capture enveloping range, few drive elements, simple control, rapid capture, etc.

An aerospace hardware derelict item is salvaged in space, modified, and reused to provide manned facilities in orbit. The hardware packages added onto the salvaged discarded item enhances its value, and incorporates at least the subsystems required to effectively reuse portions of previously discarded launch vehicle components, and other derelict objects in space. The hardware, and technique used reduces the cost of launching comparable hardware to orbit, because of the reuse, and provides a human habitation in orbit. The salvaged items include the external tank of the space shuttle, other derelict orbital hardware, the add-on cargo pod in two forms, and the ability to convert the derelict into a cost effective reusable item. The salvaged hardware is initially capable of contributing mass, length, interior volume, strongback, rotational stability mass, interior pressurized volumes, artificial gravity, and stability with the addition of simple subsystems for salvage, and interior development packages.