51 results about "CubeSat" patented technology

The invention provides an LCOS linearity correction method based on a cubesat laser communication terminal, and relates to the technical field of communication, and the method comprises the steps: obtaining the corresponding relation between the driving voltage and phase response of an LCOS device; generating an initial correction curve; performing least square fitting according to driving voltage segments, and dynamically adjusting segment points until the slope difference of boundary points of adjacent segment curves is smaller than a preset threshold value; establishing a mapping relation table between the target phase value and the driving voltage; and temperature parameters are collected in real time for compensation and correction. According to the invention, the phase modulation precision of the LCOS is improved, and the on-orbit working stability of the system is enhanced.

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.

1. Name of the Design Product: Satellite Teaching Aid (CubeSat 3U). 2. Use of the Design Product: This product is used as a teaching aid for simulating payload simulation, thermal control simulation, optical simulation, camera transmission, etc. of real aerospace satellites. 3. Design Key Points of the Design Product: Lies in the shape. 4. Picture or Photo that Best Illustrates the Design Key Points: Perspective View 1.

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.

Dual mode engine for propelling spacecraft, including combustion chamber having flange end, open nozzle end, and enclosed chamber portion extending therebetween, propellant tank in fluidic communication with combustion chamber, electronic controller, power source operationally connected to electronic controller, and fluid flow motivator operationally connected to electronic controller and connected in fluidic communication with propellant tank. Engine has chemical propulsion portion with propellant inlet port operationally connected to combustion chamber and disposed adjacent flange end, ignition trigger electrode positioned in combustion chamber adjacent propellant inlet port and operationally connected to electronic controller and operationally connected to power source propellant inlet port fluidically connected to tank electric propulsion portion with two electrodes ionizing propellant positioned in combustion chamber adjacent nozzle end, plurality of attitude control thrusters operationally connected to electronic controller and in fluidic communication with propellant tank, and plurality of valves, each fluidically connected between attitude control thruster and propellant tank.

A four element folded slot-based multiple-input-multiple-output (MIMO) antenna for use with CubeSats is described. The MIMO antenna includes a dielectric circuit board, and a metallic layer that covers a top side of the dielectric circuit board and covers a first portion and a second portion of a bottom side. A first meandering slot wraps from the top side, over the first edge and into the first portion on the bottom side. A second meandering slot wraps from the top side, over the second edge and into the second portion of the bottom side. A first feed horn is directed towards the first meandering slot. A second feed horn is directed towards the second meandering slot. A first capacitor is connected to the metallic layer across a slot section of the first meandering slot. A second capacitor connected to the metallic layer across a slot section of the second meandering slot.

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 belongs to the technical field of structural batteries, and particularly discloses a structural energy storage integrated lithium ion battery, a preparation method thereof and application of the structural energy storage integrated lithium ion battery in a cubesat. The method comprises the following steps: coating the surface of a carbon fiber braided fabric with electrode slurry to respectively obtain a battery positive electrode and a battery negative electrode; and then placing the electrolyte, the diaphragm, the battery positive electrode and the battery negative electrode in the resin matrix, arranging the carbon fiber braided fabrics on two sides, and integrally curing with the resin matrix through a vacuum infusion or lamination curing molding process to obtain the structural energy storage integrated lithium ion battery. According to the structural energy storage integrated lithium ion battery prepared by the invention, the active material of the lithium ion battery can be directly integrated into the carbon fiber reinforced resin matrix to form a composite material with structural strength and an energy storage function; and the method is applied to the fields of miniature cubesat and other spacecrafts or unmanned aerial vehicles with high requirements on weight reduction and volume utilization rate.

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.

The invention relates to a novel reliability test system which is used for evaluating the performance of a mortise and tenon connection structure used by a cubesat combination, and the test system realizes efficient and accurate performance evaluation through a mortise and tenon buckle splicing test structure and an overall test structure and by using a vibration sensor technology. A traditional tenon-and-mortise connection structure test generally uses a static stretching test method, and actual use conditions are difficult to simulate. Under the ground simulation service condition, vibration sensor signals are collected in real time, and it is ensured that the test result is closer to the actual performance. The modular structure can easily adapt to mortise and tenon joints of different sizes and types, so that the test process is highly customizable. The vibration sensor technology plays a key role in the invention and allows real-time monitoring of the vibration response of the connection structure. By recording and analyzing the vibration data, potential problems, such as fatigue cracks or signs of failure, can be detected. This contributes to early recognition of potential reliability issues, thereby improving structural quality and safety. In conclusion, the invention provides an innovative tenon-and-mortise structure reliability test system, the test process is more accurate due to the application of the vibration sensor technology, and the reliability of the tenon-and-mortise connection structure is expected to be improved in multiple fields including furniture manufacturing, constructional engineering and other application fields.

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 invention relates to the technical field of cubesat injection ejection equipment, in particular to cubesat controllable ejection equipment which comprises storage equipment, switching equipment, ejection equipment and extraction equipment, sliding grooves are formed in the two sides of the upper end face of the storage equipment, a rectangular pushing plate is arranged in the storage equipment, a first spring is arranged below the pushing plate, and a second spring is arranged below the first spring. A first driving part is arranged in the storage device, a first driving shaft is arranged at the output end of the first driving part, one end of the first driving shaft is connected with one side of the switching device, a second driving part is arranged in one end of the storage device, a second driving shaft is arranged at the output end of the second driving part, and a connecting plate is arranged on one side of the switching device. After the upper buckle of the extraction equipment loses limitation on the cubesat, the upper buckle still moves inwards by a certain distance, and the situation that release of the cubesat is interfered, and the cubesat is damaged is prevented.

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 present invention proposes a ground experiment method for the sun-oriented orientation of a cube satellite's solar wing based on gravity vector tracking. By introducing a constant gravity vector that can be analogized to the solar vector as the vector tracking source, without the need for complex dedicated equipment such as solar simulators, and only using general test equipment such as a three-axis turntable, the simulation and test of the sun-oriented orientation function of the solar wing can be achieved. This method has fewer test steps, simple procedures, high speed, does not require complex modeling, is suitable for the testing of solar wings and ground experiments in general laboratory environments, is used to verify its sun-oriented orientation function and accuracy, and is more in line with the characteristics of cube satellite research and development, such as fast speed, simple structure, and low cost.

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.

In a method for routing communications through a constellation (100) of a plurality of CubeSats orbiting a planet and responsive to a ground controller, the ground controller determines optimal tunnels (210) that include inter-satellite links (122) between CubeSats by: generating virtual nodes (110); generating the sub-satellite points (120) to correspond to points on a path through each virtual node that CubeSats follow; calculating an inter-satellite link from a sub-satellite point another sub-satellite point during a period from a time of initialization when the first sub-satellite point is at an entry point of the first virtual node until the first sub-satellite point is at an exit point; and generating an information packet describing the optimal tunnel (210) including inter-satellite links (122) that interconnect the sub-satellite points (120). The information packet transmitted from the ground controller to one of the CubeSats and is forwarded from one of the CubeSats to remaining CubeSats.

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.