1087results about "Cosmonautic component separation" patented technology

A satellite stack includes satellites arranged in a stack, where each satellite in the stack has a planar portion with an associated support structure and a subsystems housing containing satellite subsystems extending along an edge of the planar portion. For each satellite, a height of the subsystems housing is greater than a combined height of the planar portion and the associated support structure. Each satellite is oriented at a nonzero rotational angle with respect to the satellite immediately beneath, such that no satellite has its subsystems housing positioned directly above the subsystems housing of the satellite immediately beneath.

This application presents ideas for devices that solve energy generation problems and space exploration. This application show idea that is using device such using conductor / element or device. And this application show idea that is using device such using conductive line or carrier device. This application show idea that is using device such using photon or particle related device.

The invention relates to a connecting rod configuration separation mechanism for pressing and releasing a stacked satellite assembly, and belongs to the technical field of spaceflight separation mechanisms, the connecting rod configuration separation mechanism comprises a stacking device and pressing separation devices, the stacking device is installed on the corner of each satellite of the stacked satellite assembly, each pressing separation device comprises a pressing mechanism and a rotating mechanism, and the rotating mechanism is installed on the stacking device. The pressing mechanisms are installed on the stacking device of the top satellite, the rotating mechanism is installed on the stacking device of the bottom satellite, the satellites are sequentially stacked through the stacking devices on the corners to form a stacked satellite combination, each pressing mechanism comprises two pressing rods, the pressing mechanisms are connected with the rotating mechanism through the two pressing rods, and the rotating mechanism is connected with the rotating mechanism through the two pressing rods. According to the mechanism, reliable compression connection of a plurality of stacked satellites can be achieved, the action of the mechanism is released from constraint during separation, separation of the stacked satellites is achieved, and the mechanism has the advantages of being simple in structure, large in bearing capacity, small in separation impact, large in separation margin and the like.

The invention belongs to the technical field of space docking, and particularly relates to a petal type docking interface driving device for space on-orbit assembly. The butt joint interface driving device comprises a driving module, a fixing module, a locking module, an active interface and a passive interface. The locking module is in contact connection with the driving module, and the driving module and the active interface are connected with the fixing module through bolts to jointly form an active butt joint module. When the two modules are in butt joint, the petal type active and passive butt joint interfaces of different structures are connected with each other, and a locking push rod in the locking module is driven by a motor to push a locking sleeve and a lock pin to reach and leave a lock groove, so that active locking and unlocking are achieved. The butt joint interface driving device is simple and compact in structure, small in overall size, light in weight, high in reliability and suitable for rapid butt joint and separation between a space modular robot and a mechanical arm, and power-off self-locking of the motor can be achieved.

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 quaternion-based rotation separation attitude adjustment method, which comprises the following steps of: aiming at attitude requirements and angular velocity requirements of satellite separation, dividing a program angle planning process into two stages: a first stage is a directional attitude adjustment stage, and forward backward pushing is carried out through a satellite separation time point to obtain an attitude quaternion corresponding to the angular velocity of 0; taking the result as a program quaternion at the attitude adjustment ending moment; the second section is a rotation starting, stabilizing and racemization section, and according to the angular speed requirement, the quaternion of a rotation starting, stabilizing rotation and racemization program is designed; after the attitude of the rocket body is adjusted to the attitude adjustment end moment program quaternion, the angular velocity is linearly accelerated from 0 to the angular velocity required by separation according to the rotation starting, stable rotation and racemization program quaternion, and the satellite separation attitude and the angular velocity both reach expected values at the separation moment. According to the method, a universal solution is provided for one-time safe separation of multiple satellites, the problem of attitude solution ambiguity in the attitude adjusting process is avoided, high universality is achieved, and the reliability of one-time safe separation of the multiple satellites is guaranteed.

The invention discloses a satellite deck coupling connection mechanism, a satellite and a satellite deck coupling connection method, belongs to the technical field of satellite structures, and particularly relates to the satellite deck coupling connection mechanism and the satellite deck coupling connection method. The problems of low assembly efficiency, uncontrollable thermal stress and difficulty in on-orbit maintenance in the existing bolt connection technology are solved; the mechanism comprises interlocking units which are arranged on the butt joint edges of the two cabin plates in pairs. The satellite cabin plate coupling connection mechanism, the satellite and the satellite cabin plate coupling connection method are applied to modular cabin plates of spacecrafts such as the satellite and the like, and bolt-free interlocking connection aims at overcoming the defects of traditional bolt connection in the aspects of rapid disassembly and assembly, thermal deformation adaptability and the like.

The present invention relates to a ring-synchronized separation adapter for separating a payload in a space environment, and more specifically, to a ring-synchronized separation adapter for separating a payload in space environment, in which the ring-synchronized separation adapter has a configuration in which a plurality of segmented clamp portions are simultaneously fastened or unfastened by rotating a drive member making contact with all of the plurality of segmented clamp portions, so that accuracy in separation of the payload is improved, explosion is not utilized, and stability of the ring-synchronized separation adapter is improved as a whole.

An antenna array has a plurality of square or rectangular antenna assemblies. Each assembly includes a first antenna assembly surface with a solar cell and a second antenna assembly with one or more antenna elements. The antenna assemblies are interconnected without gaps therebetween to form a first contiguous array surface comprised of the first antenna assembly surfaces and a second contiguous array surface comprised of the second antenna assembly surfaces. The antenna assemblies are connected together by mechanically stored-energy connectors, such as spring tape, that self-deploy the array in space without the use of electric energy.

A satellite capture module (1) comprises: a main structure (10) extended about a main axis (X); a deployment system (40) configured to move said main structure (10); a plurality of arms (20) mounted on said main structure (10) around said main axis (X); a plurality of anchoring members (30) configured to engage an anchoring structure (103) of a target satellite (100), each anchoring member (30) being mounted on a respective arm (20) of said plurality of arms (20); wherein each arm (20) is movable between a first configuration in which said anchoring member (30) is placed at a first radial distance from said main axis (X) and a second configuration in which said anchoring member (30) is placed at a second radial distance from said main axis (X) greater than said first radial distance.

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 invention provides an explosive bolt collecting device, an aerospace craft separation system and an aerospace craft, the collecting device is installed on a separation body of the aerospace craft, the collecting device is provided with a collecting hole, the collecting hole is provided with a collecting opening, and the collecting hole extends inwards from the collecting opening in the first direction; wherein the sectional area of the collecting hole is gradually reduced in the first direction. When the separated body of the aerospace craft is separated, the explosive bolt is detonated, the broken connecting piece flying out of the explosive bolt at a high speed moves in the collecting hole in the first direction, the broken connecting piece of the explosive bolt extrudes the inner wall face of the collecting hole to form self-locking, the broken connecting piece cannot rebound, and the risk that normal work of other equipment is affected is reduced; and meanwhile, the risk that the broken connecting piece rebounds back to the mounting hole of the explosive bolt is reduced, the separation system is high in working reliability, and smooth and safe separation of the separated body of the aerospace craft is facilitated.

This invention patent discloses a wall-mounted ring unlocking device for picosatellites, specifically related to the technical field of spacecraft structures. It includes a base, a locking and releasing mechanism, a pressing ring, a thrust mechanism, and a docking ring. The base is fixedly connected to the satellite bottom plate. The locking and releasing mechanism is used for locking and releasing the pressing ring. The pressing ring is an unclosed ring structure, and both ends of the pressing ring are connected to the locking and releasing mechanism. The thrust mechanism is used to push the docking ring to separate in a direction away from the base. The technical solution of this invention solves the problems of the existing picosatellite separation mechanism, such as complex structure, heavy weight, and large unlocking impact, and enables the safe release of picosatellites.

The invention relates to the technical field of space optics, in particular to a low-impact separation device between an optical camera and a satellite platform, which comprises a controller and at least six separation supports, and the at least six separation supports are uniformly arranged at intervals in the direction of the outer envelope circumference of the satellite platform and the optical camera. The at least six separation supports are divided into a plurality of separation support sets, and the separation supports of each separation support set are adjacent to the separation supports of at least one other separation support set. The separation support comprises two supports which are connected and symmetrically arranged, one support is connected to the optical camera, and the other support is connected to the satellite. The support comprises a shell, a support butt joint end and an explosive bolt. The explosive bolt is connected to the support butt joint end and located in a containing space defined by the support butt joint end and the shell. And the controller is used for receiving a detonation signal to activate the explosive bolts, so that the explosive bolts of different separation support groups are detonated at different time, and the explosive bolts of the same separation support group are detonated simultaneously.

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.

The utility model provides a spacecraft space rendezvous and docking structure and a spacecraft, and relates to the technical field of spacecraft parts, the spacecraft space rendezvous and docking structure comprises a first docking piece, a second docking piece, a guide structure and a buffer structure, the first docking piece is used for being installed on a spacecraft body, and the second docking piece is used for being installed on the spacecraft body. The second butt joint piece is used for being installed on a body of another spacecraft. The first butt-joint piece is provided with a butt-joint part, the guide structure is arranged on the edge of the butt-joint part, and when the second butt-joint piece is close to the first butt-joint piece, the guide structure abuts against the edge of the second butt-joint piece so that the second butt-joint piece can move along the guide structure; the buffer structure is installed on the butt joint part, and when the second butt joint piece is close to the first butt joint piece, the buffer structure abuts against the second butt joint piece and is used for being pressed and shrunk. According to the spacecraft space rendezvous and docking structure, the guiding function and the buffering function can be achieved through the guiding structure and the buffering structure respectively, and the stability of a spacecraft in the space interaction docking process is effectively guaranteed.

Stacked satellite dispensing systems are described herein. The disclosed systems can release individual satellites or small batches of satellites from a satellite stack in a controlled manner. Dispensing individual satellites at different locations in space can more evenly dispense satellites or to dispense satellites at distinct locations. A deployment system, including a motor, a cable, and spacers, can allow for the individual satellites to be deployed from the satellite stack. The disclosed satellite dispensing systems also include spacers that are positioned between two or more satellites to help attenuate the force of the mass of satellites stacked above.

A gap maintaining type pyrotechnically actuated separation and unlocking device. A front load bearing cylinder (2) is connected to a front flange (1), a rear load bearing cylinder (5) is connected to a rear flange (13), a combustion chamber (11) is connected to the rear flange (13), and an igniter (14) is mounted on an end face of the combustion chamber (11); a support tube (6) is in clearance fit with an inner hole of the rear flange (13); a two-way clamping ring (4) and a one-way clamping ring (3) are fitted over the support tube (6) and serve as load bearing members for supporting the front load bearing cylinder (2) and the rear load bearing cylinder (5), and the one-way clamping ring (3) is cooperatively attached to a bevel of the front load bearing cylinder (2); a positioning ring (16) is connected to a connection tube (15), and simultaneously, the one-way clamping ring (3) is pressed, thereby implementing pre-tightening between the one-way clamping ring (3) and the front load bearing cylinder (2) and pre-tightening between the two-way clamping ring (4) and the rear load bearing cylinder (5); a connection base (10) is connected to the combustion chamber (11), an inner wall of the connection tube (15) is in clearance fit with the connection base (10), and the connection tube (15) and the support tube (6) are located on an inner side of the combustion chamber (11) and are in clearance fit with each other; and a piston disc (9) is connected to the support tube (6), and the piston disc (9) is in clearance fit with an outer diameter of a long tube in the combustion chamber (11). A pyrotechnic system, comprising the gap maintaining type pyrotechnically actuated separation and unlocking device. In this way, the device has the characteristics of high bearing capacity, small unlocking impact and having no separation impulse, and thus is suitable for connection and separation of satellite-satellite and satellite-rocket.

Container for collecting space debris. Space debris ingress through container's openable access port. A barrier deployment mechanism has a rotatable roll of flexible material on a frame assembly. A barrier translation mechanism translates the frame assembly towards or away from the access port to adjust the size of the compartment. When unwound, the roll forms a compartment for containing the space debris within the interior of the container, thus forming an obstruction between the compartment and access port.

The invention discloses a built-in fairing pre-tightening and rotary separation hinge mechanism, relates to the technical field of carrier rockets, and solves the problems that a fairing separation mode is complex in structure, an exposed separation mechanism affects the aerodynamic configuration of an aircraft, and an axial pre-tightening function does not exist. An upper fairing is movably connected with a cabin body; a hinge structure is arranged on the upper fairing, and a cabin body groove structure is arranged on the cabin body; one end of the hinge structure is inserted into an insertion groove in the cabin groove structure, the other end of the hinge structure is provided with a pre-tightening structure, the pre-tightening structure abuts against the inner wall of the cabin groove structure, and the hinge structure, the pre-tightening structure and the cabin groove are matched to achieve circumferential positioning and axial pre-tightening of the upper fairing; the cabin body is provided with a step structure, and the upper fairing is buckled on the step structure. According to the device, the circumferential positioning and axial pre-tightening functions of the fairing can be achieved, the hinge is used as a turnover fulcrum to ensure that the turnover track is controllable, and the situation of collision or dead locking is avoided; and meanwhile, the influence of the separation mechanism on the aerodynamic configuration of the aircraft is reduced.

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 gripper consists of a support structure (1) with two counter-directional carts (2, 3) moving on the support structure along linear guides (4). The carts (2, 3) are driven by a screw gear in the form of a drive shaft (5) which is counter-threaded on both sides, and of two nuts (6, 7) connected with the carts (2, 3). Each cart (2, 3) consists of two linear finger guides (8) situated at the same angle (α) relative to the axis (9) of the support structure guides (4). On the guides (8) are mounted are retractable fingers (10) which extend from the carts (2, 3) through the action of the springs (12). The flat central section of the support structure (1), parallel to the axis (9) of the guides (4) of the support structure (1), and the fingers (10) of the carts (2, 3) delineate the bottom (13) and two sides (14) of the grasping envelope (15) of the satellite, respectively. The entrance into the grasping envelope (15) is situated opposite to its bottom (13) and is equipped with optical sensors (19) detecting the presence of the adapter (11) of the satellite being gripped. Each finger (10) has a distal end (16) and a proximal end (17) relative to the entrance into the grasping envelope (15). The two sides (14) of the grasping envelope (15) are delineated by the gripping surfaces (14') of the fingers (10) which are substantialy perpendicular to the bottom (13) of the grasping envelope (15). The gripping surface (14') of each finger (10) in the vicinity of the proximal end (17) of the finger (10) has a protrusioin (18) directed towards the grasping envelope (15).

The invention discloses a stacked multi-star release mechanism adopting modular components, which is characterized by comprising a base (1), a separation seat (2), a T-shaped pressing rod (3), a pressing rod control block (4), a large pressing rod (5), a stacking column (6), a pressing cover (7), a hinge (8), a pressure spring A (9), a push rod component (10) and a pressing screw (11). The separation seat (2), the hinge (8) and the push rod assembly (10) are all modular assemblies, are simple in structure and easy to assemble and produce, and can be compatible with stacking columns in various forms.

Stackable satellites with a large diameter antenna that are configured to stack within each other like fully overlapping shallow bowls and to fit tightly within the area in a pay load fairing. The bowl-shaped satellites allow the satellites to nest within each other so that many satellites can be stacked within the height limits of the launch vehicle pay load fairing. A cradle is used to hold the stack, and tensions cables keeps the satellites in compression during launch. The satellite is an antenna with the bus integrated into the antenna structure. The antenna occupies the entire diameter / footprint of the satellite and thus can be matched up with the cross section of the fairing. This allows the largest possible antenna size to be launched without a complex deployment mechanism.

A Hold Down Release Mechanism, HDRM, interface for attachment of a spacecraft to an adjacent structure of a launch vehicle or another spacecraft, wherein the HDRM interface is configured for forming part of a single or multi-point releasable attachment of the spacecraft to said adjacent structure. The HDRM interface includes first and second connector parts, wherein one of the first and second connector parts is fastened to said adjacent structure, and other is fastened to the spacecraft; wherein the first connector part has a tapered projection with a non-circular external surface; wherein the second connector part has a tapered recess with a non-circular interior surface configured for form-lockingly receiving the tapered projection, for enabling transfer of torsion and shear load between the first and second connector parts, when the tapered projection is inserted in the tapered recess and the first and second connector parts are pressed together.

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.

To solve the problem that there is a case in which it is difficult to recover a flying object.SOLUTION: A recovery device has an engine part which burns supplied propellant to obtain a thrust in a predetermined direction, and is configured to be mounted on an airframe of a recovery object such that the recovery device can obtain, with the engine part, the thrust in a direction opposite to a direction in which the airframe of the recovery object obtains a thrust.SELECTED DRAWING: Figure 1