37 results about "CubeSat" patented technology

This invention relates to a 2U space dust sampling cubesat and sampling method. The 2U space dust sampling cubesat includes a 2U shell structure, a transparent fixed baffle, a cover plate driving unit, a camera, a lens, an aerogel collection material, a light source, an attitude adjustment component, and an onboard component. The transparent fixed baffle is disposed inside the 2U shell structure and is arranged perpendicular to the length direction of the 2U shell structure. The transparent fixed baffle divides the internal space of the 2U shell structure into a first assembly space and a second assembly space arranged front and rear. The aerogel collection material, attitude adjustment component, and light source are installed in the first assembly space. The cover plate driving unit, camera, lens, and onboard component are installed in the second assembly space. The lens is located in front of the camera and is arranged directly opposite the transparent fixed baffle. The aerogel collection material is located behind the light source and is arranged directly opposite the transparent fixed baffle and lens. A telescopic cover plate is provided on the upper side of the 2U shell structure at a position corresponding to the first assembly space.

The present invention discloses a novel micro-cathode arc thruster and an acceleration method thereof, relating to the field of CubeSat propulsion. The micro-cathode arc thruster comprises a thruster, a trumpet-shaped accelerating electrode, and a power processing unit. The high-potential output end of the power processing unit is respectively connected to the anode of the thruster and the trumpet-shaped accelerating electrode, and the cathode of the thruster is connected to the low-potential output end of the power processing unit. The power processing unit is used to generate a transient high voltage between the anode and cathode of the thruster, forming a plasma flow ejected from the nozzle of the thruster. The trumpet-shaped accelerating electrode is arranged at the nozzle of the thruster and is used to accelerate the plasma flow. The present invention achieves acceleration of ions in the plasma flow ejected by the thruster without adding an additional power device.

The terrestrial-based satellite telephone monitoring system may be used for, among other things, intelligence gathering purposes that intercept communications to or from target satellite handsets or terminals. The exemplary signal processing units receive wireless signals and extend the line-of-sight range of a terrestrial-based satellite telephone monitoring system. The exemplary signal processing unit may be installed in or on an aerial vehicle, such as an unmanned aerial vehicle (UAV), or a satellite, such as a CubeSat, or other vehicle.

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.

This disclosure relates generally to a modular CubeSat dispenser with novel features that enable universal compatibility within a single CubeSat dispenser system. The apparatus of the invention is particularly useful for 10 CubeSat dispensers and other standardized deployment devices.

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 relates to the technical field of spaceflight, in particular to a cubesat attitude control support assembly and a cubesat attitude control system.The cubesat attitude control support assembly comprises a first support and a second support, and the first support comprises a first base plate and a first mounting plate perpendicularly fixed to the first base plate; the first support comprises a first substrate, the second support comprises a second substrate parallel to the first substrate, the first mounting plate is fixedly connected with the second support so as to form an attitude control assembly accommodating space between the first substrate and the second substrate, and the first substrate, the first mounting plate and the second substrate are all used for mounting attitude control assemblies. The invention aims to provide a cubesat attitude control bracket assembly and a cubesat attitude control system aiming at at least one technical problem related in the background technology.

An improved FMEA method for CubeSats based on the Analytic Hierarchy Process (AHP) is proposed. This method involves assembling an FMEA expert team based on the overall mission and subsystem characteristics of the CubeSat; performing subsystem-level FMEA analysis; and using the AHP to calculate the subjective risk factors (SOD) for each device failure mode, combining expert experience information to calculate expert subjective weights ω = [ω1 ω2 ω3]. The method then corrects the judgment matrix provided by multiple experts and the calculated subjective weight values ​​using a tangent space objective average weighting, resulting in a comprehensive risk factor for each device failure mode, a statistical functional failure mode risk factor, and a subsystem-level risk factor. Finally, the FMEA failure modes for the entire CubeSat are calculated. This invention, considering the characteristics of CubeSats, employs a quantitative and objective expert database and the AHP to comprehensively process various component-level and functional failure modes. This approach enables rapid identification of potentially risky components and analysis of their consequences, while simultaneously providing scientific FMEA analysis results for both subsystems and the entire CubeSat.

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.

The invention discloses a wide-view-field X-ray polarization detection cubesat device which comprises a mechanical supporting structure, an X-ray polarization detector, an electronics system and a mechanical supporting structure, the X-ray polarization detector and the electronics system are located in the mechanical supporting structure, and the electronics system is used for providing a power source for the X-ray polarization detector. The X-ray polarization detector is responsible for carrying out necessary processing and conversion on signals received from the X-ray polarization detector so as to facilitate subsequent data analysis; the X-ray polarization detector is used for executing measurement of X-ray polarization; according to the X-ray polarization detector, a working mechanism of a photoelectric effect is adopted, based on a microstructure gas detector, a high-granularity gas microchannel plate GMCP is used as an electron multiplier, a charge-sensitive Topmetal pixel chip is matched to image a photoelectron track, the X-ray polarization detector can capture a fine track image of photoelectrons, and the X-ray polarization detector is used for detecting the photoelectron track. Therefore, under the condition of limited photon number, the sensitivity of X-ray polarization detection is improved to the greatest extent.

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.

The present invention relates to large-scale CubeSat chassis and methods of making and using same. The disclosed CubeSat comprises specific structural features that can yield up to at least a 27 U CubeSat that has comparable or even better performance and benefits than that of smaller CubeSats. Such volume increase coupled with the structural integrity of Applicants' CubeSats opens up a wide range of CubeSat performance increases including, but not limited to, larger payloads, more diverse payloads and larger CubeSat fuel capacities. Such increased payload capabilities includes, for example, increased focal length capabilities and such improved maneuverability results in, for example, decreased ground sample distance, increased optical resolution, increased ability to change orbits, increased structural stability. In addition, the judicious selection of such specific structural features can result increased payload capabilities and improved maneuverability.

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 invention discloses a cubesat group separation device, which relates to the technical field of spaceflight and comprises a cubesat cabin platform, a cubesat cabin and a cubesat cabin cover, wherein the cubesat cabin platform is arranged on the upper stage of the carrier rocket; the lower end of the cubesat cabin is arranged on the cubesat cabin platform; after the cubesat is placed in the cubesat cabin, the upper end of the cubesat cabin is connected with the cubesat cabin cover, and the cubesat cabin cover is used for applying pressing force to the cubesat. Cubesat launching can be completed in a large scale at a time, so that the launching frequency is reduced, and the launching risk and cost are reduced; and meanwhile, the networking period of the cubesat can be shortened, and the networking efficiency is improved.

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.