131 results about "CubeSat" patented technology

The invention discloses a separate-orbit CubeSat launching device for carrying or launching a single or a plurality of standard CubeSat. The separate-orbit CubeSat launching device mainly comprises a rectangular framework, an unlocking mechanism and a spring mechanism, wherein the rectangular framework is mainly constructed by connecting a cabin door, a baffle plate, bottom plates, side plates, upper cover plates and a back cover plate in a threaded way; a separate guide rail is arranged as a satellite slide rail; a rotary electromagnet and the like are taken as the unlocking mechanism; the spring mechanism comprises a push plate and pressure springs. The separate-orbit CubeSat launching device has high generality, and is low in cost; the transportation safety and launching reliability of satellites can be ensured, displacement of the satellites during transportation can be prevented, and the cabin door is not collided with a running platform or reversed in an opening process; moreover, a push plate anti-separation measure and the like are adopted.

This system and method provides for a plurality of satellite ground stations, distributed across some geographic region, and for these regions in turn to be scalable to cover large regions or even the globe using a combination of low-orbit satellites, terrestrial participant devices, and cloud-based communications. The invention in its simplest form is intended to solve the short temporal window problem inherent to the scenario where a single base or ground station is trying to track and communicate with a low-end LEOSAT or even a cube-satellite.

The invention discloses a CubeSat satellite platform based on a mobile phone. The CubeSat satellite platform comprises the smartphone, an interface switching module and an external module, wherein the external module comprises one or more of an attitude control module, a communication module, a maneuvering module and an extendable loading module, the smartphone is used for performing interaction with the attitude control module, the communication module, the maneuvering module and/or the extendable loading module respectively through the interface switching module to complete satellite platform tasks, the attitude control module is used for achieving attitude control of a CubeSat, the a CubeSat is used for achieving information interaction between the CubeSat and a ground station or other spacecrafts, the maneuvering module is used for controlling the CubeSat to operate in a set orbit, and the extendable loading module is used for extending load and/or adding corresponding functional units so as to achieve extendibility of the satellite platform under the control of the smartphone. The coupling problem of satellite design, tasks and progress is effectively solved, the period of study and development is short, and expansibility and reconfigurability are high.

The invention discloses a cubesat brake sail de-orbit device. The cubesat brake sail de-orbit device is composed of two completely identical cubesat brake sail de-orbit sub-devices. Each cubesat brake sail de-orbit sub-device comprises a de-orbit device body and a partition board arranged on the top of the de-orbit device body, wherein the de-orbit device body is of a central symmetry structure and comprises a main frame, an upper end cover, a sail storage chamber guide rail, a Hall sensor, a base plate and two expanding mechanisms, the main frame is Z-shaped, the main frame is divided into two identical chambers with the center of the main frame as the symmetry center, and the two expanding mechanisms are arranged in the two chambers respectively. Four film sails are expanded in four directions by means of tape spring masts to increase the normal sectional area of satellite movement, so as to successfully solve the problem that a cubesat stays on the original track for a long time after fulfilling a task and becomes space debris.

The invention discloses a multiunit cubesat main load-bearing structure. A main frame is of a rectangular frame structure and provided with rib plates matched with the shape of the long-edge section of the main frame in the direction of the long edges. End covers are arranged at the two ends of the main frame in the direction of the long edges, and the profile of the end covers is matched with the shape enclosed by the main frame. A middle skeleton is arranged at the positions corresponding to the rib plates in the main frame, and the profile of the middle skeleton is matched with the shape enclosed by the rib plates. Four threaded rods are arranged in the main frame in the direction of the long edges, penetrate through the middle skeleton and are fixed to the end covers at the two ends of the main frame. According to the multiunit cubic-star main load-bearing structure, due to the fact that the main frame is of an integrated structure, parallelism among guide rails can be guaranteed, the main load-bearing structure cannot be clamped by a launcher in the launching process, and successful launch is guaranteed; in the cubesat assembling process, internal PCBs can be assembled with the end covers, the middle skeleton, the threaded rods and nuts firstly, and then the assembled whole body is arranged in the main frame in a sleeving mode; the difficulty of the assembling process is greatly lowered, and the internal layout integrity is improved.

The invention discloses a cubesat integrated power supply system comprising a solar cell array, an energy management unit, a voltage stabilization unit, a power distribution unit, a MCU controller, anenergy storage cell unit, a separation switch circuit, and a general electric interface unit; compared with the prior art, the MPPT function is integrated to the solar energy transition unit; a separation trigger power-up unit is fused in the power supply system; the electric interface employs a general standard PC104 slot; a low power single-chip microcomputer can gather and detect system states, and can realize information interaction with a spaceborne computer, thus optimizing the power supply system integral functions, and improving the system energy transition efficiency and reliability.

The invention discloses a multi-CubeSat combination structure and a variation method thereof, and mainly relates to the technical field of CubeSat combination connection. CubeSat has multiple surfaces, and any two surfaces may need to be connected according to different combination structures; and according to the multi-CubeSat combination structure and the variation method thereof, a mode that the direction of the power supply current of an electromagnet is controlled, occupied space is smaller, and any surface of the CubeSat can achieve dead lock connection is provided, and the purpose of connection of the CubeSat which is provided with limited control units and limited installation space is achieved. According to the multi-CubeSat combination structure, any two pieces of CubeSat can be connected mainly through a simple current control system, same-type different-structure device layout and variation of the polarity of the electromagnet, a male-female type design scheme is avoided, and the multi-CubeSat combination structure has the characteristics of small size, fewer actuators, low energy consumption, high redundancy rate, and the like.

The invention relates to a networking wireless sun sensor system based on autonomous power supply. Three to six wireless sun sensors are installed out of a CubeSat. Each sensor comprises a light shield, a bottom shell and an interior circuit board, wherein the interior circuit board is installed in the bottom shell, two sides of the light shield and the bottom shell are provided with screw holes, and the light shield and the bottom shell are installed together via the screw holes and fixed on the surface of the CubeSat. The invention also relates to a working method of the networking wireless sun sensor system based on autonomous power supply. The wireless sun sensors capable of realizing autonomous power supply and a magnetometer will overcome the problems of complex installation of a conventional wired sun sensor and inaccurate measurement of a magnetic field; and the characteristics of autonomous power supply and wireless data transmission enable installation of the sun sensors and the magnetometer on satellite surface to be more flexible and freer.

The present invention discloses a Cubesat on-board computer, comprising: a voting system, a processing system, a power supply module and a memory unit. The voting system is a voter adopting MSP430 as a system. The processing system comprises Cyclone V, and a CPU 1 and a CPU 2 on the Cyclone V, wherein the CPU 1 is responsible for acquiring data, scheduling tasks, computing and sending control commands, and storing satellite data and scientific data in SD cards, and the CPU 2 only simultaneously takes actions with the CPU 1 when computing the control commands so as to provide reference for the computing results of the CPU 1. According to the Cubesat on-board computer, provided by the present invention, when the reliability of the computer is ensured, the demands of processing capability are satisfied, the integration level of the system of the satellite-borne computer and the flexibility of the design are improved, and the volume, the mass and the overall power consumption of the system are greatly reduced.

The invention provides a CubeSat catapult, comprising: a shell structure including a chamber body, a chamber cover and a combination plate, wherein the combination plate can divide the chamber body into a plurality of catapulting channels, and a catapulting guide rail is arranged in the chamber body; a plurality of catapulting structures, wherein each catapulting structure includes a spiral springand a catapulting carrier plate, the catapulting carrier plates are used for bearing a CubeSat and can slide along the catapulting guide rail, and the section of each spiral spring is in polygonal shape; and a locking structure arranged on the shell structure and used for locking or unlocking the chamber cover and the chamber body, wherein after the locking structure unlocks the chamber cover andthe chamber body and the chamber cover opens to a preset angle, the spiral springs drive the catapulting carrier plates to slide along the catapulting guide rail to catapult the CubeSat. The CubeSatcatapult is capable of catapulting CubeSats of various sizes at the premise of avoiding damage to the CubeSats and is more universal. The invention also provides a catapulting method of CubeSats.

A thruster includes multiple segments of electrically-operated propellant, electrodes for igniting one or a few of the electrically-operated propellant segments at a time, and a propellant feeder for moving further propellant segments into engagement with the electrodes. The segments may be configured to provide equal increments of thrust, or different amounts of thrust. The segments may each include an electrically-operated propellant material surrounded by a sealing material, so as to keep the propellant material away from moisture and other contaminants (and/or the vacuum of space) before each individual segment is to be used. The thruster may be included in any of a variety of flight vehicles, for example in a small satellite such as a CubeSat satellite, for instance having a volume of about 1 liter, and a mass of no more than about 1.33 kg.

The invention discloses a 2U cubesat structure which comprises side face frames, crossbeams and connecting screws. Each side face frame is composed of a rectangular truss and a reinforcing rib plate,and the crossbeams and the connecting screws are connected. The 2U cubesat structure is composed of two 1U structures, different layers are distributed in each 1U structure, and different function units can be installed on the different layers. The 2U cubesat structure is a standardized design structure, the function units can be installed in the standardized structure, through the modularized design structure, the structure is simple and easy to install, under the condition that the structural strength meets the requirement, a new material is adopted, mass is small, and cost is low.

The invention relates to a heat design method of a cubesat. The heat design method comprises the following steps of external heat flow calculation: the periodic average values of solar radiation heat flow, earthlight heat flow and earth radiation flow are calculated; heat radiation mode selection: according to the solar radiation heat flow, the earthlight heat flow and the earth radiation flow, the corresponding space heat sink temperature on each surface of a satellite and on the satellite skin can be calculated, in addition, the maximum space heat sink temperature and the minimum space heat sink temperature of each surface are determined, and the heat radiation surface is selected according to the space heat sink temperature obtained through calculation; heat radiation surface area calculation: the heat radiation surface area is determined according to the highest work temperature and the greatest space heat sink temperature required by satellite single-machine equipment; and active temperature control power determination. The heat design method has the advantages that parameters such as the heat radiation area and the active control power are accurately calculated, the problems of many parameters (for example, the heat radiation area and the active control power are set according to the experience of engineers), control equation nonlinearization and calculation iteration complexity of the traditional heat design analysis method are solved, and the development period is greatly shortened.

The basic idea of this invention is to send one or more cubesats into orbit, each equipped with a hardware security module. Users would send their transaction to the cubesats which would collect them into blocks, sign them, and send (bounce) them back to earth (and to one another). Bounce Blockchain provides scalability through sharding (transactions will be partitioned over cubesats). Because modern hardware security modules are tamper-resistant (become inoperable if tampered with) or tamper-responsive (erase their keys if tampered with), take their keys from physical processes, and have been validated, socio-technical protocols can ensure that it is infeasible to forge the identity of a hardware security module in a cubesat with another cubesat. If, however, some cubesats are destroyed, the blockchain will continue to execute correctly though some transactions will be lost. New cubesats can be sent up in short order as they are quite cheap to launch. If, in spite of these assurances, some cubesats fail traitorously, the blockchain can survive through algorithms similar to Practical Byzantine Fault Tolerance techniques.

The invention discloses a hinge for an expandable solar cell array of a cubesat. The hinge for the expandable solar cell array of the cubesat comprises a body installation plate fixing piece, an expendable array fixing piece, a positioning shaft and a torsional spring. The body installation plate fixing piece is of a three-side surrounding structure. The expandable array fixing piece is matched with the shaped defined by the body installation plate fixing piece. The positioning shaft is fixed in the body installation plate fixing piece and is in running fit with the expandable array fixing piece. The axis of the positioning shaft serves as the circle center around which the body installation plate fixing piece and the expandable array fixing piece rotate mutually. Limiting protrusions which are located on the same rotating surface are arranged on the opposite end faces of the body installation plate fixing piece and the expandable array fixing piece in a staggered mode. The body installation plate fixing piece and the expandable array fixing piece are in clearance fit through the limiting protrusions on the end faces of the body installation plate fixing piece and the expandable array fixing piece. The positioning shaft in the expandable array fixing piece is sleeved with the torsional spring and limited in the axial direction by the two inner end faces of the expandable array fixing piece. A straight torsional arm of the torsional spring makes contact with a body installation plate and the expandable array. The hinge is small in size, low in weight, high in adaptability and high in transportability.

The invention discloses a cubesat two-dimensional centroid adjusting method. The method comprises the following steps of step one, adjusting an upper surface of a horizontal table board to horizontalthrough a horizontal adjusting mechanism; step two, establishing a satellite body coordinate system and a triaxial accelerometer body coordinate system; step three, placing a triaxial accelerometer measurement module on the horizontal table board, and calibrating a triaxial accelerometer; step four, suspending the triaxial accelerometer measurement module, and adjusting the centroid of the measurement module to a center of form location; step five, installing the triaxial accelerometer measurement module on the top of the cubesat; step six, suspending the cubesat through a suspending mechanism, and reading a measurement value of the triaxial accelerometer measurement module in the premise that the satellite is stable, and judging a centroid offset direction and an offset location size; andstep seven, adjusting the cubesat to the geometric longitudinal axis location through the centroid adjusting module. The centroid adjusting method disclosed by the invention is high in precision, good in practicability and convenience.

The present invention discloses a power source bus-bar circuit applied to the CubeSat. A first P- MOS transistor Q1 and a second P- MOS transistor Q2 are connected in parallel; the S pole of the first P- MOS transistor Q1 and the S pole of the second P- MOS transistor Q2 are connected together, and then are connected with one end of a second resistor R2, the S pole of a third P- MOS transistor Q3 and one end of a first resistor R1; the G pole of the first P- MOS transistor Q1 and the G pole of the second P- MOS transistor Q2 are connected together and then are connected with the other end of the second resistor R2, one end of a third resistor R3, one end of a fourth resistor R4, one end of a fifth resistor R5, the D pole of the third P- MOS transistor Q3 and one end of a second ceramic capacitor C2; and the D pole of the first P- MOS transistor Q1 and the D pole of the second P- MOS transistor Q2 are connected together and then are connected with one end of a seventh resistor R7, the positive pole end of a third tantalum capacitor C3 and the positive pole end of a fourth tantalum capacitor C4. The power source bus-bar circuit of the invention has the advantages of small size, high functional integration and small energy loss. With the power source bus-bar circuit adopted, the requirements of de-energization of the CubeSat after a ground test is performed and the independent energization of the CubeSat after the CubeSat enters an orbit can be satisfied.

An agile attitude control system (AACS) that is a three axis attitude control device for small spacecraft based on miniature single gimbal control moment gyroscopes (SGCMGs) actuators. The AACS enables agile attitude slewing and accurate pointing/tracking for spacecraft made of multiple CubeSat units, or, more generally, for nanosatellites.

The invention discloses a ground education satellite assembly capable of achieving a standard CubeSat function and a demonstration method of the ground education satellite assembly. The assembly comprises a unit satellite body, a ground demonstration terminal and a Lora wireless communication module used for achieving interaction between the unit satellite body and the ground demonstration terminal, wherein the demonstration terminal sends uplink control commands, and a satellite selects different actions to be executed and downward returns data information after receiving the different commands; and a user adjusts and changes a satellite operation condition and acquires data information collected by the satellite via the corresponding uplink command. The ground education satellite assembly can achieve tests on functions of a standard CubeSat, comprising air pressure measurement, communication, positioning, image acquisition, satellite posture measurement and adjustment and the like, also can be applied to design, assembly, integration and test of the satellite, software programming, hardware system development and satellite ground measurement and control, and is applicable to training of personnel at all levels and learning and debugging aiming at the CubeSat. The ground education satellite assembly has the advantages of being compact in structure, complete in system, comprehensive and intuitive in demonstration function and the like, and is applicable to education of aerospace engineering majors.

The invention discloses a cubesat release device. The device comprises a satellite cabin, a first spring, a push plate, cubesats, a pressing block, a pin shaft, a sliding block, a sliding rod, a second spring, a moving frame, a push rod, a third spring, a sliding lead screw, a motor and a stop rod. A plurality of cubesats are fixed in the satellite cabin and the to-be-released cubesat is limited by utilizing a pressing plate; when the cubesat is released, a moving frame and the pressing plate are driven by the motor to release the cubesats; and meanwhile, the cubesat released in the second syn-position is limited by the stop rod, a release device with a cubesat intermittent release function is achieved, the structure is stable, and the intermittent release function can be achieved.

The invention relates to a manned spacecraft on-orbit service and maintenance verification method. The method comprises the following steps of a building a test environment, transporting materials required by on-orbit verification to an extraterrestrial space station, and quickly building the test environment by utilizing the existing resources of a manned aerospace system and the flexible controlcapability of an astronaut; b assembling materials in a cubesat module form in a space station cabin, and carrying out on-orbit comprehensive detection on the materials; and c releasing the target inthe materials out of the space station cabin, and performing an approaching observation test and a net capturing test on the target. Verification is carried out on the manned spacecraft on-orbit service and maintenance technology, and compared with other technical verification methods, the method has the advantages of being good in technical basis, high in supporting capacity, small in task risk,large in number of guarantee resources, low in test cost, high in efficiency, comprehensive in verification and the like.