575results about "Cosmonautic component separation" patented technology

The invention discloses a super-large plane array type space computing power center, and belongs to the field of space technologies and space infrastructures. The space computing power center is deployed in a morning and evening orbit and comprises a plurality of computing power satellites which are physically docked in orbit through a docking mechanism. Each computing power satellite comprises a satellite platform, and a computing power load, a condensation type energy system and a space pump drive fluid loop system which are arranged on the satellite platform, and a docking mechanism for inter-satellite in-orbit physical docking is also assembled on the satellite platform. A plurality of computing power satellites form a planar array type rigid whole body through the docking mechanism, and energy scheduling among the light condensation type energy systems is realized based on the docking mechanism. According to the space computing power center, through multi-satellite physical docking and energy collaborative scheduling, the energy utilization efficiency and the power supply reliability of the system are improved.

The invention discloses a separating and unlocking mechanism. The separating and unlocking mechanism comprises a base; the separating and pushing assembly comprises a fixed shell and a push rod, the fixed shell is detachably arranged on the base, and the push rod penetrates through the fixed shell and is slidably arranged on the fixed shell; the locking piece is detachably connected with the push rod and is positioned outside the fixed shell; the connection unlocking assembly comprises a guide rod and a clamping jaw structure, the guide rod is connected with the end, deviating from the locking piece, of the push rod, the separation unlocking mechanism adopts a connecting rod linkage assembly to control the locking piece to unlock the push rod, single-point faults are avoided, and reliability is improved; the separation pushing assembly and the connection unlocking assembly form an independent unlocking module, the structure is compact, the space is saved, multi-stage unlocking control is achieved, the unlocking and locking stability is improved, the operation convenience is improved, unlocking lagging is avoided, and the requirement for unlocking force is reduced.

The locking and releasing device for the spacecraft comprises a base, a shell, a sliding sleeve, a pin puller, a split nut assembly and a first spring, the base is of a cylindrical structure, and a first mounting plate and a second mounting plate are fixed to the two axial ends of the base correspondingly; the sliding sleeve slidably sleeves the base and is connected with the first mounting plate through the first spring. The shell sleeves the sliding sleeve at an interval and is fixed to the first mounting plate and the second mounting plate. The pin puller is mounted on the shell, and one end of the pin puller extends into the shell; a plurality of notches evenly distributed in the circumferential direction are formed in the side wall of the base, a first through hole communicated with the interior of the base and used for ejecting the locking rod is formed in the middle of the first mounting plate, a split nut assembly used for locking the locking rod is mounted in the base, and a plurality of sets of connecting lug plates evenly distributed in the circumferential direction are fixed to the outer side wall of the split nut assembly. The multiple sets of connecting lug plates are arranged in one-to-one correspondence with the multiple notches and penetrate through the notches through connecting rods to be hinged to the sliding sleeve.

End effectors and systems may capture, release, and / or create a mating engagement between the end effector and a target object, such as in capture and launch applications, grasping applications, or pick-and-place applications. The end effectors are tolerant of positional and rotational misalignment of the target object, and may include a plurality of roller wheels, one or more of which is arranged in a non-parallel plane with respect to one or more other roller wheels. Control devices enable operation of these and other end effectors in a portable, self-contained system that a single operator can use in a hand-held or wearable form factor. The control devices may include a drive system for power transfer, manual and automatic leveling or angular positioning mechanisms, and / or operator-mounting features. Related methods may involve capture and release, pick-and-place, and / or grasping operations performed with the disclosed control device and one or more end effectors coupled thereto.

A tiered spacecraft docking station is adapted to facilitate docking of spacecraft within outer space. A first tier includes a first frame enclosing a first area. A first net-like mesh is coupled to the first frame and fills the first area enclosed by the first frame. A second tier includes a second frame enclosing a second area. A second net-like mesh is coupled to the second frame and fills the second area enclosed by the second frame. A plurality of support beams attach the first frame to the second frame. A lander is used to slow and stop a spacecraft on a celestial body. The lander includes a first and a second webbed structure and a decelerator coupled to the first webbed structure and / or the second webbed structure. The decelerator maintains a tension in the first and / or second webbed structure below a predetermined threshold.

The invention relates to a large-bearing memory alloy driven unlocking separation nut mechanism, and belongs to the technical field of spaceflight structures. The memory alloy separation nut comprises a shell assembly, a split nut assembly, a bearing force transmission assembly, a memory alloy driving assembly and a nut separation assembly, a modular design concept is adopted, configuration design is carried out on the memory alloy separation nut according to three functional parts of driving, transmission and execution, the bearing efficiency and the unlocking reliability are improved in a targeted mode, and the memory alloy separation nut is suitable for being used for a large-scale industrial production. The unlocking impact is reduced, and the separation nut is more convenient to use.

Device for separating a second part from a first part, including a first interface intended to be fixed to the first part and a second interface intended to be fixed to the second part, a docking stud passing through the first and second interfaces and being held at the first interface by a separation nut and at the second interface by a first nut / locknut assembly, the docking stud being covered by a cap secured to the second interface and intended to receive it once the separation nut has broken and the first and second interfaces have been separated, and a retraction spring is provided to extract the docking stud from the first interface, propel it towards a damping material fixed at the bottom of the cap and hold it there by its residual holding force alone.

The invention relates to the technical field of lunar cabin section assembly and butt joint, and discloses a large-deformation flexible inflation channel for lunar cabin section assembly and butt joint, which comprises a sealed flexible channel and at least two groups of deformation control air bags, the two ends of the sealed flexible channel are respectively connected with a butt joint frame and a butt joint ring, and the deformation control air bags are arranged on the outer side of the sealed flexible channel; the deformation control air bag controls the overall pose and rigidity of the sealed flexible channel through inflation and deflation so as to compensate pose deviation between the butt joint cabin sections and balance internal loads generated by butt joint, the large-deformation flexible inflation channel is assembled on the lunar surface cabin sections in a butt joint mode, and the large-deformation flexible inflation channel can be folded and unfolded on the lunar surface in a lossless, efficient and repeated mode. Multiple properties such as wear resistance and puncture resistance and good air tightness are achieved, and multiple times of connection, separation and reconstruction of the cabin section can be achieved; the air bag can be controlled through deformation to balance the internal pressure load, prevent the cabin section from overturning and the like according to the posture change and the load; and the method also has high reliability and long-term service capability, and adapts to pose changes caused by lunar soil settlement.

A launch system and method for orbital or suborbital air-launch of a payload involving releasably coupling a launch vehicle with a towed aircraft via an articulatable carriage to form an air-launch assembly, towing the air-launch assembly via a tow aircraft and interconnected tow cable to a first altitude, releasing the air-launch assembly from tow at or above the first altitude, activating the towed aircraft propulsion system and initiating a pull-up and climb maneuver of the towed aircraft to a second altitude, articulating the articulatable carriage to shift the air-launch assembly from a stowed position to a deployed position with the launch vehicle spaced from the towed aircraft, releasing the launch vehicle from the articulatable carriage and thus from the towed aircraft, and activating the launch vehicle propulsion system for further altitude gain or to meet specific mission requirements.

A method includes providing an image to a feature extraction model to generate feature data. The image depicts a portion of a first device and a portion of a second device. The feature data includes coordinates representing key points of each of the first and second devices depicted in the image. The method also includes obtaining position data indicating a position in 3D space of a connector of the first device. The method further includes providing the feature data and the position data to a trained autonomous agent to generate a proposed maneuver to mate the connector with a connector of the second device.

A connector (4, 4') with a load transfer rod (40, 40') designed to transfer forces and / or moments from a first support body (42) connected or connectable to the load transfer rod (40, 40') to a second support body (44) connected or connectable to the load transfer rod (40, 40') is characterized in that the load transfer rod (40, 40') has a first rod element (41) and at least a second rod element (43) which are connected to each other in a joining section (45, 48) of the load transfer rod (40, 40') by means of a joining device (46), wherein the joining device (46) has a lower tensile and / or compressive strength than the respective rod element (41, 43) above a predetermined temperature threshold.

The invention discloses a motor-driven steel ball locking mechanism based on gear and thread pair linkage, and belongs to the technical field of spacecraft on-orbit docking. The mechanism comprises a motor serving as a power source, wherein an output shaft of the motor is fixedly connected with a cylindrical gear; the cylindrical gear is in meshed connection with a locking mechanism through a gear; a butt joint conical rod is arranged above the locking mechanism, and the locking mechanism is used for capturing and locking the butt joint conical rod. When a guide taper sleeve in the locking mechanism makes contact with the butt joint taper rod, a contact switch is triggered, a motor is started, a thread pair is driven to move through gear transmission, a locking screw is made to move axially, and a steel ball is controlled to lock or unlock the butt joint taper rod. The contact switch is triggered through contact between the guide taper sleeve and the butt joint taper rod, space dust and radiation interference are avoided, autonomous capture is achieved, and the butt joint precision is remarkably improved. The mechanism is highly integrated and modularized and can be adjusted according to requirements of different spacecrafts, the size and weight of the spacecrafts are reduced, and the utilization efficiency of spaceflight resources is improved.

A propulsion system 1 with three propulsion points of a space module 10 for orbital deployment for at least one satellite, said propulsion system 1 comprising a chassis 2 having exactly three first housings 6 each formed to receive a propulsion unit 3 and at least one second housing 7 formed to receive a tank 4, at least one liquid fuel tank 4 arranged in the second housing 7, and exactly three propulsion units 3, each propulsion unit 3 arranged in one of the three first housings 6 and having at least one thruster 30.

This invention relates to a satellite panel coupling connection mechanism, a satellite, and a satellite panel coupling connection method, belonging to the field of satellite structural technology. Specifically, it addresses the problems of low assembly efficiency, uncontrollable thermal stress, and difficult on-orbit maintenance inherent in existing bolted connection technologies. The mechanism includes paired interlocking units located at the mating edges of two panels. This satellite panel coupling connection mechanism, satellite, and satellite panel coupling connection method are applied to modular panels in spacecraft such as satellites, aiming to overcome the shortcomings of traditional bolted connections in terms of rapid assembly / disassembly and thermal deformation adaptability through boltless interlocking connections.

This invention provides a satellite-launch separation signal system and a satellite-launch separation system, comprising: a first user unit that outputs two satellite-launch separation signal sources, each leading to two different contacts of a first analog separation relay; these signals then return to the first user unit via a first separation connector and a second separation connector; the input and output terminals of the first analog separation relay are each connected to a satellite display socket via cables; and a second user unit that outputs two satellite-launch separation signal sources, each leading to two different contacts of a second analog separation relay; these signals then return to the second user unit via a first separation connector and a second separation connector; the input and output terminals of the second analog separation relay are each connected to a satellite display socket via cables. This invention ensures that the signal is provided by the analog satellite-launch separation relay during ground testing, and avoids fatal malfunctions such as accidental separation commands issued before launch or mis-separation due to severe vibrations during the active phase, thus improving the reliability of the satellite-launch separation signal link.

A reusable orbital vehicle (14) for a space transport system (10) comprises a fuselage (22) the shape of which widens in a direction extending from the front end (24) of the reusable orbital vehicle to a rear end (26) of said fuselage (22) and inside which there is defined a housing (32), and a crew-evacuation vehicle (34) housed in said housing (32) and oriented in such a way as to be extracted from the housing (32) in a direction (D1) directed from said rear end (26) of the fuselage (22) toward said front end (24) of the reusable orbital vehicle (14).

The application relates to the technical field of spacecrafts, and discloses a four-redundancy pyrotechnic mechanism lock which comprises a lock body assembly and a pressing assembly which cooperate with each other, the lock body assembly comprises an igniter, an ignition shell, a pin rod, a seat body, a structural upper plate and a hinge module, the pressing assembly comprises a ball seat, a pressing rod, a pull rod module and a structural lower plate, the hinge module is arranged in the inner cavity of the seat body, the ignition shell is arranged on the seat body and oppositely forms two groups of ignition cylinder bodies, two groups of pin rods are arranged in each of the ignition cylinder bodies respectively, the pin rods movably cooperate with the hinge module, the side surface of the ignition cylinder body is provided with the igniter which communicates with the inner cavity, the ball seat is arranged on the structural lower plate, the one end of the pressing rod is provided with the pull rod module, the other end of the pressing rod penetrates through the ball seat and movably cooperates with the hinge module. The four-set unlocking execution mechanisms are completely independent, any unlocking execution mechanism can realize unlocking when normally working, so that the working reliability of the pyrotechnic mechanism lock is improved.

The invention provides a gas-driven hook lock type locking and separating mechanism for satellite-rocket separation, and belongs to the technical field of spaceflight. Air cylinder driving is adopted to be matched with a three-connecting-rod transmission mechanism, and unlocking of the non-initiating explosive device is achieved. Compared with explosive bolts and other initiating explosive devices, the risk that extremely high instantaneous impact load generated by explosion threatens precise instruments on satellites and the risk that the surface of the satellites is polluted by fuel gas and chippings can be avoided; and meanwhile, pneumatic driving and mechanical transmission are adopted, so that the mechanism can be repeatedly locked and separated. The device comprises a rocket end connecting base and a passive end connector, the rocket end connecting base is installed at the tail stage of a rocket, and the passive end connector is connected with a satellite load through bolts; the locking claws are symmetrically mounted on the rocket end connecting seat through hinge shafts, and locking grooves matched with the passive end connectors are formed in the locking claws; the device further comprises a driving device and a transmission connecting rod mechanism, and the transmission connecting rod mechanism is connected between the driving device and the locking clamping jaw.

A propulsion-based payload transporter (PPT) and methods of its operation are provided. The PPT may operate in a gravitational field such as that on the lunar surface. Via rocket engines that produce thrust, the PPT is configured to descend onto and over a payload to be moved. Once in position with respect to the payload, latches engage with the payload so as to secure the payload in a payload holding bay of the PPT. The rocket engines produce thrust with an exhaust plume that is angled away from the payload to avoid payload damage. The rocket engines, which are located at the sides of the payload holding bay, produce thrust to propel the PPT and the payload in a substantially horizontal trajectory with respect to gravity and toward a landing destination.

The invention relates to a resupply space station (10) for resupplying a spacecraft in orbit around a celestial body with propellant, the space station comprising a main body (11) carrying a store of resupply propellant (16). A flexible fluid pipe (18) that is capable of extending and bending outside the main body (11) connects this store (16) to a connector (19) for fluidic connection to the spacecraft carried by a nanosatellite (12). This nanosatellite (12) comprises a propulsion module, a docking sensor, a system (21) for reversibly anchoring to the spacecraft and a system for transferring fluid from the store (16) to the spacecraft via the pipe (18) operable when the nanosatellite (12) is anchored to the spacecraft.

Angular velocity control device for a spacecraft and corresponding spacecraft. The invention relates to a device (1) for controlling the angular velocity of a decommissioned spacecraft, comprising: - a stator (3) and a rotor (4) movable about an axis (A21) of rotation relative to the stator. The stator (3) comprises an electrically conductive and non-ferromagnetic body (6), while the rotor (4) comprises a magnetized system (7) configured to induce, in the stator (3), braking eddy currents of a relative motion of the rotor (4) with respect to the stator (3); - a levitation magnet (11) intended to cooperate with a magnetic field generated by an external source to induce magnetic levitation of the rotor (4) with respect to the stator (3). The device (1) is made of one or more non-ferromagnetic materials in a zone (11ZI) influenced by the magnetic field generated by the levitation magnet. SAMPLE FIGURE: [Fig.4]

The present application relates to a kind of locking release mechanism, including locking release connection assembly, with the locking release connection assembly and the locking jaw assembly of cooperation locking, and the axial drive assembly for driving jaw assembly to open or close;Locking release connection assembly includes locking lever, locking lever sequentially passes through first connecting plate and second connecting plate and is locked with jaw assembly;Jaw assembly includes two locking jaws, support frame, pin shaft, any locking jaw is hinged to support frame by pin shaft, any locking jaw is provided with torsion spring, and support frame is fixed on second connecting plate;Axial drive assembly includes drive rod, drive rod moves along the axis, and drives two locking jaws to open or close;The end of locking lever is provided with the clamping groove matched with two locking jaws.The present application can improve the stability of the connection of jaw assembly and locking release connection assembly, no excess during separation release process, no pollution, small impact, simple and reliable.

The invention relates to a satellite capturing and locking device, and belongs to the field of satellite docking. The satellite capturing and locking device comprises an active end and a passive end, the driving end comprises a driving end base, a guide mechanism, a driving mechanism and two locking mechanisms; the guide mechanism comprises a guide frame and a guide plate; the guide plate is in an inverted cone shape; the passive end comprises a passive end base and a caught rod; the guide plate is used for guiding the captured rod into the opening; the driving mechanism is used for driving the two locking mechanisms to move oppositely so as to clamp the captured rod, and therefore locking between the driven end and the driving end is achieved. The caught rod at the driven end is clamped left and right through the two locking mechanisms to achieve locking, the two locking mechanisms are arranged in a bilateral symmetry mode, namely, the two locking mechanisms are arranged at intervals in the transverse direction, the axial distance between two satellites after locking is reduced, and therefore the size of a satellite platform is reduced. The locking mechanism comprises a first-stage locking unit and a second-stage locking unit, and the ejection speed during separation can be guaranteed.

The invention relates to the area of space science, in particular, to the area of industrialization of outer space and, directly, to the facilities for inserting and moving various objects in circumplanetary cosmic space. It is intended to solve geocosmic problems in industrial-scale volumes - to carry out research and development, special-purpose, touristic and other types of works and services in outer space and stabilize global climate. Artificial orbital ring complex by Yunitski is made in the form of a ring satellite (1), located in a circular orbit (5) around a natural cosmic body (2) in the equatorial (4) plane (3), which comprises: structural frame (7), utilities (8) and communications (9), power units (10), transition galleries (11) positioned in annular housing (12) with protective shell (13), gate chambers (14) with docking units (15), configured to dock with space vehicles (16), residential (19) and research-production (20) units, equipped with systems of control, life support, active and passive (17) protection against the harmful effects of space (18). The structural frame (7) is longitudinally tensioned, and the ring satellite (1) is in circular motion in orbit around the natural cosmic body (2) with velocity V0, m / sec, defined by the ratio: 1≤V0 / V1H≤1.01, where V1H, m / sec, - first space velocity for the equatorial circular orbit of the location of the ring line connecting the centers of mass of the cross sections of the artificial orbital ring complex, at altitude of H0, m, defined by the ratio: 0.02≤H0 / R0≤0.5, where R0, m, - radius of natural cosmic body (2) in equatorial (4) plane (3).

The satellite supporting cabin comprises an upper end frame, a truncated cone shell and a lower end annular flange which are integrally formed by carbon fiber composite materials, the upper end frame is formed by wrapping a first foam sandwich with the carbon fiber composite materials and solidifying the first foam sandwich, and the upper end frame is used for being connected with a satellite support. The lower end annular flange is used for being connected with a substage of an aerospace vehicle. The satellite supporting cabin can achieve the purpose of reducing the weight and has high rigidity.

This application relates to the field of aerospace technology, and in particular to an irregularly shaped multi-satellite distributor. The lower end face of its irregularly shaped load-bearing structure is abutted against the upper end face of its conical load-bearing structure. The lower end face of the irregularly shaped load-bearing structure is circular, and the upper end face is elliptical. Its load-bearing beam is fixed within the mounting cavity of the irregularly shaped load-bearing structure. A top satellite mounting bracket is fixed to the upper end of the irregularly shaped load-bearing structure, and a side satellite mounting bracket is fixed to the outer side of the irregularly shaped load-bearing structure, perpendicular to the major axis direction of the upper end face. An irregularly shaped satellite mounting bracket is fixed to the outer side of the irregularly shaped load-bearing structure, perpendicular to the minor axis direction of the upper end face. Regularly shaped satellites are fixedly mounted on the top satellite mounting bracket and / or the side satellite mounting bracket, and irregularly shaped satellites are fixedly mounted on an irregularly shaped satellite mounting bracket. This application allows the protruding points of the satellites to be enveloped within a safe envelope range, and avoids increasing the size of the fairing.

The invention discloses a thermal unlocking structure for a spaceflight satellite. The thermal unlocking structure comprises a first structural part, a second structural part, a thermal unlocking assembly, a winding assembly and a restraint strap. After the restraint strap is fused by the hot knife, the knife holder retreats under the action of the spring, and the stop pin synchronously unlocks the limiting piece on the winding roller; the winding roller is driven by the pre-tightening torsion spring to rotate rapidly, and the fused restraint strap is completely rewound into the closed cavity in the shell through the guide wire. According to the design, it is ensured that restraint strap residues generated in the separation process are actively stored and restrained, and the risk that the restraint strap residues freely drift away to become space junk or interfere with a satellite body and an unfolding mechanism is fundamentally eliminated. The whole recovery process does not need extra energy, only depends on elastic potential energy pre-stored by the mechanism for driving, has the advantages of high reliability, low impact and no secondary pollution, and remarkably improves the safety and cleanliness of an on-orbit unlocking and releasing task of the spacecraft.

This provides a coupling mechanism that can reliably protect the internal mechanism from adverse effects of the external environment. [Solution] The coupling portion 3A of one robot module 2A comprises an annular base portion 4 having an opening in the center, a planar flexible sealing member 4 covering the opening of the base portion 4, an engaging portion 7 disposed outside the sealing member 4, and a drive unit 9 built inside the base portion 4 separated by the sealing member 4. The coupling portion 3B of the other robot module 2B comprises an engaged portion 7' ​​that engages with the engaging portion 7. The drive unit 9 engages the engaging portion 7 via the sealing member 4, and the sealing member 4 deforms in sync with the movement of the engaging portion 7.