25 results about "CubeSat" patented technology

The application discloses a kind of inductively coupled plasma micro-propeller of double antenna structure, and its discharge system includes gas source, air inlet valve, antenna, discharge tube and radio frequency power supply.Gas source is directly connected with discharge tube by valve, wherein the valve can control the air intake during discharge;Then two identical helical antennas are connected in series at a certain distance and are wound on the same discharge tube, the output end of radio frequency power supply is connected to the end of antenna away from air inlet through impedance matcher, and the other end of radio frequency power supply and antenna is connected to the same ground.The application has simple structure, small space size and mass, and only needs to work stably at low power and low air intake, realizing micro-newton to millinewton level thrust.It is suitable for equipping in most of the current constellation micro-satellites without propulsion system of nanosatellite and cubic satellite.In addition, the length and spacing of the two antennas in the system are adjustable, and the size of the propeller can be adjusted according to the specific requirements of micro-satellite.

The invention discloses a small space debris-oriented passive capturing and disposing spacecraft, cubesat and method, solves the problems that in the prior art, small space debris cannot be captured, and an existing capturing device performs active capturing, and has the beneficial effects that the space debris is passively captured and thoroughly eliminated. According to the specific scheme, the passive grabbing and handling spacecraft for the small space debris comprises a chip satellite and a capturing assembly, a plurality of solar sails are connected to the peripheral side of the chip satellite, the chip satellite controls the bending state of the solar sails so as to control flight of the spacecraft, and the capturing assembly comprises a folding storage wall plate, a triggering platform and a bistable capturing arm; the folding storage wall plate is connected to the peripheral side of the chip satellite through the hinge, the hinge is a bent hinge and has elasticity, the triggering platform is connected with the chip satellite, and after the bistable capturing arm passes through the triggering platform, the bistable capturing arm is in a straight state when a spacecraft is in a non-working state.

The utility model relates to the field of spaceflight, in particular to a solar wing folding and unfolding device which comprises a supporting shell and a solar wing unfolding device, the supporting shell comprises a bottom end plate face, a peripheral plate face, a limiting vertical plate, a top end plate and a supporting column, the bottom end plate face is fixed to the bottom of a cubesat, and the peripheral plate face is hinged to the periphery of the bottom end plate face. The top end plate is connected with the peripheral plate surface through a disconnectable connecting rope, the supporting column is arranged at the center of the bottom end plate surface, the solar wing unfolding device comprises a control driving part, an extension part and a flexible solar wing, the control driving part is provided with a spring rod, a torsional spring rod and a limiting groove, the spring rod is located above the bottom end plate surface, the torsional spring rod is located above the spring rod and fixedly connected with the spring rod, and the limiting groove is formed in the bottom end plate surface. Through mutual arrangement of the spring rod, the torsion spring rod, the limiting groove and the limiting vertical plate, the purpose of automatically controlling the torsion spring rod is achieved, meanwhile, the torsion spring rod which accumulates force in advance drives the tail end rod to be unfolded in a fan-shaped mode, and the purpose of automatically unfolding the solar wing is achieved.

The utility model relates to the field of spacecrafts, in particular to an adjustable solar wing unfolding device which comprises a supporting shell and an unfolding device, the supporting shell comprises a main shell body and a penetrating opening, the main shell body is fixed outside a cubesat, the penetrating opening is formed in the peripheral face of the main shell body, and the unfolding device is arranged on the supporting shell body. The unfolding device comprises an external containing part, a solar wing part, a driving motor, a driving shaft, an adjusting rack, an adjusting gear shaft, a supporting ring, a supporting fixing ring and a main telescopic rod, the external containing part is fixed to the upper portion of the penetrating opening, and the external containing part comprises a containing shell, an extrusion spring plate, a partition plate, a solar wing plate and a connecting shaft; the solar wing plate is magnetically connected with the supporting and fixing ring through the connecting shaft, and the inclination angle of the solar wing plate is adjusted through the arrangement of the adjusting shaft and the adjusting gear shaft fixed on the solar wing.

The present disclosure provides a system for autonomously docking a spacecraft, the system comprising: one or more electromagnets; one or more sensors; and a controller. The controller configured to: identify a target; cause movement of the spacecraft closer to the target; determine a change in an orientation of the one or more electromagnets and the mating adapter on the target; and adjust a current in the one or more electromagnets to produce a torque, wherein the torque causes the spacecraft to align with the target.

The invention discloses a cubesat deploying device, which relates to the field of spaceflight and comprises a cabin body, a cabin door, an unlocking device, a cabin door buffering and locking device and a boosting separation mechanism, one end of the cabin door is rotationally connected with one side of the upper end of the cabin body; the unlocking device is installed on the outer side of the cabin body and connected with the other side of the cabin body and the other end of the cabin door. The cabin door buffering and locking device is installed on the outer side of the cabin body and connected with one side of the cabin body. The boosting separation mechanism and the cubic star are both installed in the cabin body, the lower end of the boosting separation mechanism is connected with the bottom of the cabin body, the cubic star is placed at the upper end of the boosting separation mechanism, and the boosting separation mechanism pushes the cubic star to get out of the cabin for separation. According to the application, the cabin door unfolding locking impact can be reduced, and the launching mechanical environment is optimized, so that the cubesat is effectively prevented from being damaged due to impact, and the launching reliability is improved.

A CubeSat attitude stabilization method based on PID controller parameter setting combines an inertia weight decline PSO particle swarm optimization algorithm and a variable learning rate BP neural network to carry out PID controller parameter dynamic setting, BP neural network weight parameter initial values suitable for a current CubeSat model are rapidly obtained, PID parameters are dynamically adjusted in a control process by using a BP-PID framework, and the PID parameters are dynamically adjusted in the control process. According to the dynamic parameter setting method under different external disturbance conditions, the requirements for cubesat attitude control performance such as required response time and control precision and the requirement for control robustness are met at the same time.

The utility model relates to the technical field of cubesat energy, in particular to a locking device of a multi-angle solar wing, which comprises a female chain, a male chain, a connecting shaft, a torsion spring and a control part, the end part of the female chain is symmetrically provided with a positioning hole and a positioning seat, the positioning seat on one side is provided with a fixing hole, and the positioning seat on the other side is provided with a matching hole; a limiting strip is arranged on the inner circumferential face of the matching hole, a mounting hole and a mounting base are symmetrically arranged at the end of the male chain, a mounting base hole is formed in the end face of the mounting base, a positioning ring is arranged in the mounting base hole in one side, a first limiting wheel is arranged on the end face of the mounting base on the other side, and a connecting shaft is rotationally arranged in a fixing hole; according to the utility model, the flexible unfolding of the solar wing in the range of 0-90 degrees and 0-180 degrees is realized through the slidable moving cylinder and the electromagnetic control technology.

This invention discloses a method for thermal equilibrium testing of a CubeSat, comprising: thermally modifying a prototype CubeSat; calculating the characteristic point temperatures of the modified prototype CubeSat under high and low temperature conditions; constructing a thermal equilibrium testing system and placing the modified prototype CubeSat within the system; obtaining the thermal equilibrium test boundary temperatures under high and low temperature conditions based on the characteristic point temperatures of the modified prototype CubeSat under high and low temperature conditions, respectively; and conducting a thermal equilibrium test based on the thermal equilibrium test boundary temperatures under high and low temperature conditions. This invention has the advantages of accurate testing, low cost, high reliability, short cycle time, and large-scale production.

The utility model relates to a folding and unfolding device for displaying load on a cubesat. The load is displayed; one end of the display load is a hinged end, and the other end is a free end; the hinge end is provided with a hinge piece, and the display load is mounted on a non-sailboard surface of the cubesat through the hinge piece, so that the free end has a stroke for rotating around the hinge end; when the free end rotates to be parallel to the non-sailboard surface, the display load completes folding, and the hinge piece is in a pre-tightening state; when the free end rotates to be perpendicular to the non-sailboard face, the display load is unfolded completely, and the hinge piece is in a natural state. The display load can be folded, so that in the non-display stage of spacecraft launching and in-orbit operation, space can be effectively saved, overall layout and design of a cubesat are facilitated, and the space utilization rate is high.

A set of compact microscopes, one of which works as an attachment to a camera-equipped cell phone, that can fit in a 3 U CubeSat and be used as a platform for high-resolution optical microscopy and characterization of nanoscale dynamics. An advanced software package, which can be deployed via a mobile interface, enabling users to reliably capture image sequences of moving particles with their smartphone cameras, and enabling immediate access to precise rheological results.

A hybrid propulsion system includes a housing, at least two electrodes, a solid-grain fuel material, a combustion chamber, an oxidizer port, and a nozzle. The housing has a first end and a second end and defines a cavity. The electrodes extend into the cavity. The fuel material is free of an oxidizer and is positioned in the cavity. The fuel material has a combustion surface and is exposed to the electrodes. The combustion chamber is defined between the combustion surface and the second end. The oxidizer port provides a flow of oxidizer to the combustion chamber. The nozzle is positioned at the second end. Combustion of the fuel material in the combustion chamber may be dominated by radiative heat transfer.

This application relates to the field of aerospace technology, specifically to a cubic star attitude control bracket assembly and a cubic star attitude control system. The assembly includes a first bracket and a second bracket. The first bracket includes a first substrate and a first mounting plate vertically fixed to the first substrate. The second bracket includes a second substrate disposed parallel to the first substrate. The first mounting plate is fixedly connected to the second bracket to form an attitude control component accommodating space between the first substrate and the second substrate. The first substrate, the first mounting plate, and the second substrate are all used to mount the attitude control component. The purpose of this application is to address at least one technical problem mentioned in the background art by providing a cubic star attitude control bracket assembly and a cubic star attitude control system.

A hybrid magnetometer and attitude determination and control system (hybrid magnetometer attitude determination and control system). Many scientific and commercial satellites in low-Earth orbit (LEO) use attitude determination and control systems that enable control and knowledge of the spacecraft's attitude (the orientation the spacecraft). The hybrid magnetometer and attitude determination and control system is a single package that can be integrated into small satellites, such as CubeSats, and consists of a three-axis search coil AC magnetometer and three-axis quad-mag DC magnetometer. The search coil sensors also act as torque rods for part of the orbit. Noise cancellation algorithms enable scientific grade magnetic measurements from the system without needing a sensor boom.

The application provides a global environment monitoring CubeSat system, which is composed of three non-coplanar CubeSats, the CubeSat adopts a 6u size rectangular body and a configuration of an unfolded solar wing, has a three-axis stable attitude control capability, and comprises a comprehensive information management module, an attitude control module, a power module and a data processing and transmission module. The system can monitor the concentration and source of greenhouse gases of the earth, and simultaneously monitor the ocean nuclear pollution of the earth.

This utility model relates to a multi-combination, expandable CubeSat battery base module. It includes individual batteries, with several individual batteries forming a base unit; a base frame is arranged circumferentially within the base unit; connecting parts are provided laterally and longitudinally within the base frame; and limiting components are installed within the base frame, which constrain the position of the individual battery arrays arranged within the base unit. This utility model allows for flexible assembly of the base module in a plane or three-dimensional space, adapting to the complex spatial layout inside a CubeSat and meeting the on-orbit energy storage and power supply requirements of various CubeSat models. The design of the base frame and limiting components ensures sufficient stability for the individual batteries.

The application discloses a kind of on-orbit deployment devices of CubeSat, belong to aerospace technology field.The device includes: main frame assembly, hatch, hatch unlocking mechanism, push plate assembly, lateral compression assembly and hatch compression assembly, the main frame assembly shell is square semi-enclosed body, the hatch is rotatably connected on shell by hatch unfolding locking assembly and can form closed body with shell;The hatch unlocking mechanism is installed in the left frame outside;The push plate assembly is arranged in the shell inside and is connected on the bottom plate, for pushing out CubeSat;The lateral compression assembly is arranged in the shell inside right upper corner guide rail support frame, for lateral compression CubeSat;The hatch compression assembly is fixedly installed in the inside of hatch, for CubeSat in the direction perpendicular to bottom plate is compressed.The device has high reliability, low separation angular velocity, ultra-high vacuum environment unlocking and has satellite lateral compression and the like characteristics.

The utility model relates to an auxiliary tool device adaptive to cubesat of multiple models. The device comprises a first adjusting assembly and a second adjusting assembly. The first adjusting assembly has a travel moving in the width direction; a second adjusting assembly is arranged on the first adjusting assembly, and the second adjusting assembly has a stroke moving in the length direction; through cooperative adjustment of the first adjusting assembly in the width direction and the second adjusting assembly in the length direction, the cubesats with different shapes can be adapted. The tool can meet the adaptation requirements of cubesat of multiple models, the universality of the tool is greatly improved, the structure is simple and compact, the practicability and the operation convenience are improved, and meanwhile the manufacturing cost and the maintenance difficulty are reduced.

The invention discloses a micro propeller and a preparation method thereof. The micro propeller comprises a silicon substrate provided with a heating chamber and a Laval nozzle, an outlet of the heating chamber is communicated with the Laval nozzle, and the heating chamber is used for containing ultrapure water; the heating electrode is arranged on the silicon substrate, at least part of the heating electrode is located on the heating chamber so as to make contact with the ultrapure water in the heating chamber, and the heating electrode is electrically connected with an external power source; and the glass substrate is arranged on the silicon substrate and the heating electrode. In the invention, the heating chamber and the Laval nozzle are directly integrated on the silicon substrate, and the heating electrode is arranged above the heating chamber for accommodating the ultrapure water, so that the controlled phase change evaporation of the ultrapure water is realized to generate corresponding thrust; therefore, the stable attitude control impulse can be provided on the premise that the volume mass of small satellites such as a cubesat or a nanosatellite platform is not obviously increased.

A hybrid magnetometer and attitude determination and control system (hybrid magnetometer attitude determination and control system). Many scientific and commercial satellites in low-Earth orbit (LEO) use attitude determination and control systems that enable control and knowledge of the spacecraft's attitude (the orientation the spacecraft). The hybrid magnetometer and attitude determination and control system is a single package that can be integrated into small satellites, such as CubeSats, and consists of a three-axis search coil AC magnetometer and three-axis quad-mag DC magnetometer. The search coil sensors also act as torque rods for part of the orbit. Noise cancellation algorithms enable scientific grade magnetic measurements from the system without needing a sensor boom.

A set of compact microscopes, one of which works as an attachment to a camera-equipped cell phone, that can fit in a 3U CubeSat and be used as a platform for high-resolution optical microscopy and characterization of nanoscale dynamics. An advanced software package, which can be deployed via a mobile interface, enabling users to reliably capture image sequences of moving particles with their smartphone cameras, and enabling immediate access to precise rheological results.

The embodiment of the application discloses a kind of on-orbit intelligent monitoring devices of CubeSat, it is related to satellite monitoring technical field, by utilizing the solar wing plate of existing rotation setting on the cabin plate of CubeSat, as the deployment mechanism of installation monitoring camera, without additional increase deployment mechanism, the structure of monitoring device is simplified, the control demand of deployment mechanism and the corresponding reduction of bearing use are reduced, the reliability and adaptability are relatively increased, and because the length of solar wing plate is greater than the length of connected cabin plate, when solar wing plate rotates to solar wing plate and cabin plate adhere, camera can be stored in the top end of cabin plate, avoid the overall structure of camera to protrude, affect overall stability, when solar wing plate is retracted, by the locking cooperation of first locking piece and second locking piece, so that solar wing plate and camera can be stably settled, further improve the reliability of device.

This application discloses a multi-redundancy cube satellite power supply system, which includes a solar panel and its interface module unit, a charging management unit, and a battery pack unit. The charging management unit is located between the solar panel and its interface module unit and the battery pack unit, and integrates a main MPPT circuit unit, a backup MPPT circuit unit, and a direct power supply circuit unit for the solar panel. The three circuit units are connected in parallel and serve as backups for each other, charging the battery pack unit through the aforementioned circuit units. The advantages of this technical solution are: (1) high power supply redundancy and strong reliability; (2) significantly improved energy conversion efficiency; (3) comprehensive safety protection and low operational risk; (4) high integration and strong adaptability. All circuits of the charging management unit are integrated on a dedicated circuit board, which is small in size and light in weight, and can be directly adapted to cube satellites of different specifications. The circuit design is modular, which facilitates later maintenance and upgrades.

A set of compact microscopes, one of which works as an attachment to a camera-equipped cell phone, that can fit in a 3U CubeSat and be used as a platform for high-resolution optical microscopy and characterization of nanoscale dynamics. An advanced software package, which can be deployed via a mobile interface, enabling users to reliably capture image sequences of moving particles with their smartphone cameras, and enabling immediate access to precise rheological results.

This invention discloses a method for switching attitude control modes for a CubeSat, specifically relating to the field of satellite attitude control technology. The steps are as follows: After separation of the satellite from the launch vehicle, the attitude control system is powered on, the pitch axis momentum wheel begins to rotate, and the satellite performs bias momentum angular rate damping. When the orbit is deemed valid, the filter is activated. When the filter converges and attitude is acquired, bias momentum control is performed; if a fault occurs, the system switches to angular rate damping mode. Finally, a command is injected to switch to zero momentum control mode. When the filter converges, zero momentum control is performed; if a fault occurs, the system switches to angular rate damping mode. When one momentum wheel fails, the other three wheels form either bias momentum control or zero momentum control mode to complete the predetermined attitude control task. Therefore, the above attitude control mode switching method makes full use of limited hardware resources, has clear mode switching logic, and is simple to implement in software, effectively improving the reliability of the CubeSat attitude control system.