1192results about "Cosmonautic power supply systems" patented technology

The invention belongs to the technical field of aerospace equipment, and discloses an unfolding mechanism and a spacecraft, and the unfolding mechanism is used for winding and unwinding a curling arm and comprises a bearing part, a rotating shaft and a plurality of guide parts. The rotating shaft is rotationally arranged on the bearing piece, one end of the curling arm is fixed to the rotating shaft, and the rotating shaft is used for winding and unwinding the curling arm. The multiple guiding parts are distributed on the bearing piece along the periphery of the rotating shaft and used for providing guiding in the winding and unwinding process of the winding arm. The curling arm is wound through the rotating shaft, and the curling arm can be unfolded or folded by rotating the rotating shaft, so that the solar wing adopting the unfolding mechanism is simple in overall structure, small in size and capable of meeting the space requirement of a small spacecraft. Comprising the unfolding mechanism is small in size and can meet the requirement of miniaturization design.

The invention provides a winding type flexible solar wing unfolding mechanism and a star body. The winding type flexible solar wing unfolding mechanism comprises a solar wing driving mechanism, a hinge, a rigid solar cell array, a flexible wing, a first pressing and releasing device and a second pressing and releasing device. The solar wing unfolding mechanism has the advantages of being small in folding size, light in weight, capable of being unfolded in orbit and large in folding-unfolding ratio. In addition, the structure is simple, the unfolding driving force source is simple, movement is reliable, unfolding driving kinetic energy is provided by releasing winding strain energy, and the unfolding rigidity of the solar cell array is provided after in-orbit unfolding.

The invention discloses a morning orbit computing power satellite, and belongs to the field of aerospace technology and space infrastructure. The satellite comprises a satellite platform and a computing force load arranged on the satellite platform. The satellite platform is provided with a condensation type energy system and a space pump drive fluid loop system. The condensation type energy system adopts an expandable condensation type solar cell array to generate power, switches a power supply mode according to an illumination period and an earth shadow period of a morning and night orbit, and supplies power to computing force loads and platform equipment. The space pump drive fluid loop system adopts an expandable flexible radiation cooling plate for heat dissipation, heat dissipation is conducted on computing force loads through a fluid loop, and a heating assembly is arranged to prevent a working medium from being frozen under the low-temperature working condition. Through the expandable structure and the orbit environment adaptability design, continuous energy supply and effective heat management are provided for the computing power load in the morning and night orbit environment, and high-power-consumption computing power load in-orbit stable operation is supported.

The invention discloses a modular micro-nano satellite integrated power system suitable for batch production and an assembly method thereof. The integrated power system is composed of a semi-open type electric box structure and a modular circuit board. The electricity box structure comprises an electricity box upper cover and an electricity box lower cover, and a plurality of threaded holes are formed in the two sides of the electricity box upper cover and the two sides of the electricity box lower cover and used for being connected with a satellite structure cabin plate; the heat dissipation plate bodies are distributed on the two sides of the modular circuit board; the modularized circuit board is provided with a plurality of circuit boards which are respectively provided with electronic modules in a stacking arrangement. A plurality of electronic modules are integrated into an electronic system by adopting a mechanical-electrical-thermal integrated design method, the system breaks through the function limit of a traditional satellite subsystem, the defects of information exclusion and resource waste are eliminated, a satellite is used as a whole to be subjected to unified design and resource planning tradeoff from the system level, and the system performance is improved. The weight of cables and connectors between electronic modules is greatly reduced, and the effective load ratio is improved.

The invention relates to an ultralight and ultrahigh-storage-ratio flexible solar wing, and belongs to the technical field of solar wing devices. The device comprises two solar box blankets, two scissor rods, two single-fork hinges, a pressing device and a lifting rod. Wherein the lifting rod is axially and horizontally placed; the two solar box blankets are axially and symmetrically arranged on the two sides of the lifting rod in parallel. The middle positions of the opposite side walls of the two solar box blankets are each provided with a scissor rod. A single-fork hinge is mounted at the bottom of each scissor rod; the two single-fork hinges are in butt joint with the lifting rod from the two sides of the lifting rod respectively. A pressing device is arranged on the side wall of the solar box blanket; clamping of the solar box blanket structure is achieved through the pressing device; the flexible solar wing has the advantages of being light in weight, small in folding size, good in configuration expandability and the like.

The invention relates to a scissor type unfolding mechanism which comprises a top short rod assembly, scissor type stage group rod piece units, a bottom supporting rod assembly, a driving unit assembly and an unfolding in-place indicating switch, the unfolding mechanism is connected with an unfolded object through the top short rod assembly, and the multiple stages of scissor type stage group rod piece units are connected through hinge shafts to form a multi-stage unfolding mechanism. The long rods of the multiple stages of unfolding mechanisms are connected in sequence; the bottom supporting rod assembly is connected with the driving unit assembly, the driving unit assembly drives the bottom supporting rod assembly to achieve unfolding and folding of the unfolding mechanism, after the unfolding mechanism is unfolded, a stable triangular structure is formed through the bottom supporting rod assembly, the scissor type stage group rod piece units and the limiting devices, the structural rigidity is improved, and the unfolding mechanism is controlled to be unfolded.

Provided herein are various enhancements for solar panels and photovoltaic array assemblies. In one example, a system includes solar panel assemblies each comprising a frame defining a mounting area formed from modular segments having receiving features configured to accept spline elements that mount a perimeter of a flexible photovoltaic panel into the mounting area and apply a radial tension to the flexible photovoltaic panel. A stowage mechanism is configured to hold the solar panel assemblies in a stacked configuration and comprising a rod assembly disposed through apertures in the frames of each of the solar panel assemblies. A deployment mechanism is configured to deploy the solar panel assemblies from the stacked configuration into a deployed configuration and comprising self-opening hinges attached between selected pairs of the frames.

A modular satellite system is provided comprising a main body member, one or more controllers, a communication system, a datastore, a power unit, and one or more orbital cameras. The controller(s) may be operable to perform a mission instruction. The communication system may be in communication with the controller(s) and with a client terminal. The datastore may store data accessible by the controller(s). The orbital camera(s) may be operable to capture one or more image(s) associated with the mission instruction, which may be defined as captured image(s). The controller(s) may be operable to detect and identify one or more predetermined object(s) in the one or more captured image(s). The controller(s) may generate a mission analytics packet, and the controller(s) may be operable to transmit the mission analytics packet to the client terminal.

A structural element, such as for use in conjunction with a satellite, may include one or more thin film solar cells (108-i), which are supported on a substrate (102), and a plurality of thin film antenna elements (110-i), which are supported on a carrier layer (104). The substrate (102) functions as a ground plane for the antenna elements (110-i).

The invention discloses a two-dimensional single-unfolding hundred-square-meter-level flexible solar wing which comprises a flexible array surface capable of infinitely transversely and / or longitudinally expanding based on a paper folding structure, an auxiliary unfolding mechanism for assisting the unfolding of the flexible array surface, and an unfolding driving mechanism for driving the unfolding of the whole wing, the flexible array surface is formed by connecting first folding and unfolding units located at the two ends and a second folding and unfolding unit located in the middle through piano type hinges, and the whole wing can be unfolded in the transverse direction and the longitudinal direction from the folded state through single-time unfolding. The two types of folding and unfolding units are double-vertex seven-fold paper folding units, and the flexible array plane has a single degree of freedom; the auxiliary unfolding mechanism is connected with the first type of folding and unfolding units and is connected with the unfolding driving mechanism; the motion trail of the tail end of the auxiliary unfolding mechanism in the unfolding process is consistent with the motion trail of the flexible plates on the two sides of the first folding and unfolding unit. According to the invention, a paper folding configuration is applied to flexible array surface design, two-dimensional single-time unfolding of the whole wing is realized, and the problem of insufficient storage ratio of a traditional one-dimensional unfolding configuration solar wing is solved.

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.

The invention belongs to the technical field of solar wing unfolding machinery, and particularly relates to a shears-fork type solar wing unfolding device based on a planar four-bar mechanism, which comprises an unfolding base assembly, a countersunk head assembly, a guide mechanism, a positioning assembly, a connecting assembly and a top plate arranged on the side wall of the connecting assembly, the shear fork type solar wing unfolding device based on the plane four-bar mechanism is composed of the rigid unfolding mechanism and the flexible solar wing film, and the flexible solar wing film is supported and tensioned after the shear fork type solar wing unfolding device is completely unfolded; the anti-bending section coefficient of the rod piece is high, and meanwhile, the rod piece is positioned in the axial direction in the form of an end cover and a set screw; the connecting pieces at the two ends of the shear fork rod facilitate complete positioning of pressing, the guide assembly is arranged, the folding condition of the solar wing is conveniently maintained, a guide rope keeps certain compactness and rigidity in the unfolding process of the whole mechanism, and when the flexible solar wing adopts a four-point tensioning mode, the system has higher fundamental frequency, and the stability of the system is improved. And the stability and the reliability are improved.

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 telescopic folding and unfolding mechanism for a spacecraft and the spacecraft. The telescopic folding and unfolding mechanism comprises a power source, a foundation plate, a first telescopic plate, a second telescopic plate and a third telescopic plate, wherein the foundation plate, the first telescopic plate, the second telescopic plate and the third telescopic plate are of hollow structures; one end of the power source is installed at the connecting end of the foundation plate, the other end of the power source extends out of the open end of the foundation plate, the first sleeving end of the first telescopic plate slidably sleeves the inner side of the open end, and the second sleeving end of the second telescopic plate slidably sleeves the inner side of the first telescopic end of the first telescopic plate. The third sleeving end of the third telescopic plate is slidably arranged on the inner side of the second telescopic end of the second telescopic plate in a sleeving mode. A first pulley wound with a first driving steel wire is arranged at the other end of the power source, one end of the first driving steel wire is connected to the inner wall of the connecting end of the foundation plate, and the other end of the first driving steel wire is connected to the second sleeving end of the second telescopic plate; the second telescopic plate is connected with the foundation plate through a second driving steel wire, and the first telescopic plate is connected with the third telescopic plate through a third driving steel wire.

Embodiments generally relate to small satellites. Example small satellites may include a satellite body having a rectangular tile shape, with a substantially flat front side having a length and width and being separated from a substantially parallel back side by a depth, wherein the depth is equal to or less than a third of the length and the width. A patch antenna array may be positioned on the front side and / or a large patch antenna may be positioned centrally on the front side. A solar panel may be positioned on or coupled to the back side.

The invention relates to a unilateral speed limiting type linkage device. Comprising an unfolding mechanism, a rotating shaft rope wheel assembly, a root hinge fixing rope wheel, an inter-rod hinge fixing rope wheel, an inter-rod hinge rope collecting wheel assembly, a rope collecting wheel trigger arm, a deflector rod, a small ball, a ball blocking frame, a rope blocking ring assembly, an upper stretching rod linkage rope, a lower supporting rod linkage rope, a speed limiting driving source, a speed limiting driving source output shaft and a root rope guide wheel. Linkage release of the rotating shaft rope wheel assembly can be achieved after unfolding is finished, and then the linkage state and the recovery state of the upper stretching rod linkage rope wheel are switched; the space unfolding mechanism is simple in structure and light in weight, and speed-limiting unfolding of the space unfolding mechanism can be controlled by controlling rotation of the rotating shaft rope wheel assembly through the speed-limiting driving source.

The invention discloses an energy scheduling method, system and device for a low-earth-orbit satellite and a medium. The method comprises the steps that the energy storage level is calculated according to the current generation power of the low-earth-orbit satellite, the current remaining electric quantity of an energy storage unit and the health degree of a battery; generating a future illumination time sequence according to the orbit parameters of the low-orbit satellite and the satellite attitude control data; according to the current operation state data of each device loaded on the low earth orbit satellite and a preset task list, generating corresponding task prediction energy consumption data arranged in a descending order according to the importance level through a historical energy consumption model; according to the energy storage level, the future illumination time sequence and the task prediction energy consumption data, calculating the dominant power supply amount, matching the energy consumption demand in the task prediction energy consumption data, and carrying out the preferential power supply of the equipment of the task item with the high importance level. And carrying out power supply adjustment on the other equipment by adopting at least one strategy of delaying to the illumination period, reducing the operation power and shortening the operation duration so as to realize efficient scheduling on the energy of the low-orbit satellite.

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.

A telescopic boom system provides an extensible boom that may be used in spacecraft applications including supporting photovoltaic panels, communication antennas, instrumentation, and so forth. A stowed configuration is volumetrically compact, including the boom and actuators such as a motor. During deployment, threaded nuts for each nested section of the boom are self-aligning with respect to a leadscrew driven by the motor. Sections are staged for extension in staged sequence by a flexure arm engaging a ramp feature on a portion of the nested section. Extension failure mitigation is enhanced by allowing partial retraction of some sections during extension. Once fully extended, tension of the boom may be later adjusted, modifying the structural fundamental frequency. A ratchet may be engaged with extension of a final nested section to prevent retraction of the extended boom.

The utility model relates to the technical field of solar wing unfolding mechanisms, in particular to a self-release solar wing unfolding mechanism. According to the self-releasing solar wing unfolding mechanism, the first elastic pieces in the compressed state are adopted to provide unfolding elastic force for the corresponding shear fork units; a scroll constant-force spring, a fixed pulley and a synchronous linkage assembly are adopted to link each scissor unit to unfold at a constant speed; and when the shear fork unit is unfolded to a target position, the shear fork unit and the synchronous linkage assembly are limited and locked by adopting the limiting and locking assembly. The utility model relates to a passive unfolding mechanism capable of completing self-release unfolding operation, belongs to a solar wing battery, and solves the problem that an active driving unfolding machine in the prior art is complex in structure and heavy in weight, so that the launching cost is increased.

Artificial satellites are dependent on solar arrays to produce the power needed to support the functional components of the satellite. The present invention is directed to a hinge-locking mechanism designed to maximize the photovoltaic real estate of a solar array of an artificial satellite. The hinge-locking mechanism facilitates the deployment of the solar array upon the artificial satellite entering orbit. The hinge-locking mechanism utilizes a tapered pin with an internal spring resting against an asymmetrically oval shaped cam of a common pivot point. The hinge-lock trades kinetic energy of a solar panel for spring compression to permanently lock the solar array in place upon deployment while reducing the shock load and maximizing deployed stiffness.

The invention belongs to the technical field of solar wings, and discloses a passive unfolding mechanism which comprises a bearing part, a rotating shaft, a curling arm and an abutting assembly. The rotating shaft is rotationally arranged on the bearing piece. One end of the curling arm is fixed to the rotating shaft, the curling arm can be wound to the rotating shaft, elastic potential energy is stored after the curling arm is wound, and the elastic potential energy is used for driving the rotating shaft to rotate so as to unfold the curling arm. The abutting assembly is arranged on the bearing piece, and the abutting assembly is configured to abut against the curling arm when the curling arm is unfolded, so that resistance is applied to the curling arm. According to the passive unfolding mechanism provided by the invention, the curling arm is wound through the rotating shaft, and the curling arm can be unfolded or folded by rotating the rotating shaft, so that the solar wing adopting the unfolding mechanism is simple in overall structure, small in size and capable of meeting the space requirement of a small spacecraft. The passive unfolding mechanism provided by the invention does not need to be powered by a motor, and is simple in structure and low in use cost.

An antenna assembly has a solar layer having one or more solar cells generating solar power, an antenna layer connected to the solar layer and having electronic components utilizing the solar power generated by the solar layer, and a thermal dissipation device dissipating heat locally at the antenna assembly. A large number of antenna assemblies are connected to form an antenna array in which heat is generated locally at each antenna assembly and dissipated locally at each antenna assembly.

The invention relates to the field of pressing unlocking mechanisms, in particular to a hot knife unlocking mechanism and a satellite. The heat knife unlocking mechanism comprises a rope which penetrates through the first component and the second component and is locked; the first device comprises a first shell used for being fixed with the first component; the first fixing part is arranged in the first shell and connected with the first end of the rope; the at least two hot knife assemblies are oppositely arranged, and knife bars of the hot knife assemblies abut against the two sides of the rope respectively; the second device comprises a second shell used for being fixed with a second component; and the pre-tightening adjusting assembly is connected with the second end of the rope, is arranged on the second shell and can move to adjust the tension of the rope. The locking device can be connected with a component to be locked or unlocked through the first device and the second device, locking is conducted through a rope, the pre-tightening adjusting assembly adjusts the rope to change tension, and pressing force can be changed during locking. The rope is fused through the hot knife assembly to unlock and separate all parts, impact force is small, safety is high, the structure is compact, occupied space is small, and installation and rope replacement are convenient.

The invention provides a large spaceflight load based on modular design. The large spaceflight load comprises a platform, a docking mechanism, a solar wing / antenna integrated system and a mechanical arm. The docking mechanism is arranged in the platform; the solar wing / antenna integrated system is arranged at the bottom of the platform; the two ends of the mechanical arm are connected to mechanical arm bases arranged on the platform respectively, two-dimensional rotation can be achieved with one base as the axis, and an information / power supply link between loads is established through the mechanical arm. The butt joint mechanism comprises a driving end and a driven end which are connected with each other, when the load is in a butt joint preparation state, the driving end is converted into an extension state from a contraction state, extends out of the side face of the platform and is connected with butt joint hole sites of different loads, and when butt joint is completed, a butt joint lock of the head is locked, and butt joint of the different loads is achieved. According to the load, automatic assembly can be achieved in the on-orbit assembly, replacement and maintenance process, other auxiliary devices are not needed in the assembly process, the flexibility is high, and the use benefits of the system are ensured.

The invention relates to the technical field of satellite power supplies, in particular to a weak light power-on control circuit and processing method for a satellite before satellite-rocket separation, and the circuit comprises a storage battery pack, an autonomous power-on unit, a forced shunting unit and a plurality of solar cell subarrays. The forced shunting unit comprises a forced shunting signal output module and a forced shunting power execution module, and the solar cell sub-array irradiated by weak light before satellite-rocket separation is a target solar cell sub-array; the forced shunting signal output module is arranged at two ends of the storage battery pack and is used for forming a driving voltage of a forced shunting signal through voltage division; and the forced shunting power execution modules are arranged at two ends of the target solar cell sub-array and are used for receiving the forced shunting signals to form a forced shunting loop, performing forced shunting on weak light current formed by the target solar cell sub-array before satellite-rocket separation, and pulling down the driving voltage to disconnect the forced shunting loop after satellite-rocket separation. According to the invention, the weak light current formed by the target solar cell subarray can be forcibly shunted before satellite-rocket separation.

The invention provides a solar synchronous orbit quick response SAR satellite load platform integrated configuration, which comprises a satellite body, an SAR antenna, a solar array, a phased array data transmission antenna, a star sensor and a body-mounted solar cell, the satellite body is in a quadrangular prism shape, the cross section of the satellite body is an isosceles trapezoid, the satellite body comprises a group of opposite parallel surfaces and a group of opposite inclined surfaces, the phased array data transmission antenna and the star sensor are respectively arranged on two parallel surfaces, and the SAR antenna and the solar array are respectively arranged on two inclined surfaces; the body-mounted solar cell piece and the solar array are installed on the same inclined plane and located on the inner side of the solar array, and the solar array comprises two solar array plates and has an unfolded state and a folded state. According to the method, load platform integration is taken as a design thought, so that the AIT process of the SAR satellite and the time required by AIT are greatly shortened; meanwhile, the attitude requirements of energy and load work are met, and the reliability of the initial injection stage is improved through sufficient energy.

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.

An artificial satellite and artificial satellite system has power transmission and reception efficiency and includes an expandable solar panel that generates power, expandable power transmitting antennas, and expandable beam members to connect the solar panel and the power transmitting antennas to each other. The beam members are expanded from being provided in the artificial satellite, on a plurality of rotational symmetry axes of an imaginary polyhedron formed with substantially a center of the artificial satellite as an origin. The beam members are arranged at substantially equal distances in directions of the rotational symmetry axes. A light receiving surface of the solar panel and installation surfaces of the power transmitting antennas, when expanded, are substantially perpendicular to a solar panel rotational symmetry axis passing through a center of the solar panel, and being arranged on opposite sides of a plane passing through the center of the artificial satellite.