20 results about "Spacecraft" patented technology

A tethered spacecraft is tethered to a base using a semirigid stand-off and a moving tether operated as a belt drive and operated in tension so that the tethered is precisely positioned in tension from the base spacecraft.

The invention provides a monitoring method of the internal stress of a spacecraft composite material structure in an aging process. The method includes following steps: step1, placing carbon fiber prepreg, a special optical fiber, carbon fiber prepreg, demolding cloth, and adhesive-absorbing felt on a molding mold of the spacecraft composite material structure in sequence, and performing integral sealing by employing a vacuum bag; step 2, performing curing by employing an autoclave process, performing demolding after cooling, and obtaining the spacecraft composite material structure implanted with the special optical fiber at the inner part; and step 3, placing the spacecraft composite material structure in a clean factory building, and performing aging treatment on the spacecraft composite material structure by employing a natural aging method. According to the method, based on a back Rayleigh scattering principle, the special optical fiber is implanted in the spacecraft composite material structure, the monitoring of the internal stress in the aging process is realized by employing an optical frequency domain reflection technology, the spatial resolution is high, the measuring precision is high, the repeatability is good, and basis is laid for improvement of the aging method of the spacecraft composite material structure and the improvement of the dimension stability.

The invention relates to composite material frames for spacecrafts such as satellites and the like. The problem to be solved is to provide a method for realizing a light-weight high-accuracy composite material frame, and the high-accuracy requirement on the installation of effective loads is met through processes. The device is characterized by mainly comprising rectangle cross section rod pieces[1], joints [2], joint spacers [3], joint bushing blocks [4] and rod built-in pieces [5] which are made of composite materials, wherein a frame main body consists of a certain number of rectangle cross section rod pieces [1] and joints [2] in the integral structure; the rod built-in pieces [5] are hinged in the rectangle cross section rod pieces [1]; the joint spacers [3] are hinged on the outer surfaces of the joints [2]; and the joint bushing blocks [4] are hinged inside the joints [2]. By the method, the problems of installation and supporting between the effective loads and bodies of the spacecrafts such as the satellites and the like are solved, the installation accuracy of the effective loads is effectively ensured, and the local rigidity requirement on the installation of the effective loads is met simultaneously; and the method has good effects of ensuring the machining stability of the integral structure and improving the applicability and the reliability of the structure.

The present invention relates to satellite autonomous relative navigation. In order to solve the problem of long-distance autonomous relative navigation, the present invention provides a relative navigation method using the line-of-sight angle information on the satellite. ;Establish the system model in the modified spherical coordinate system; carry out filter calculation. Compared with the existing technology, its advantages and beneficial effects are: no active ranging equipment is required to work, only the elevation angle and azimuth angle of the target need to be detected by the passive tracking and aiming equipment, combined with the relative navigation initial value, the relative navigation can be passed The filtering algorithm realizes real-time autonomous estimation of relative position and relative velocity, and provides the required information for relative orbit control. The relative navigation with only angle information is realized, so that the active star in the orbital rendezvous mission can realize autonomous relative navigation at a long distance, and the reliability of the rendezvous mission is improved.

The invention discloses a flexible spacecraft attitude and orbit tracking control method based on an expansion interference observer, relates to the technical field of aerospace, and aims to solve the problems of poor anti-interference capability and accuracy in a flexible spacecraft tracking control process in the prior art, and the method comprises the following steps: firstly, establishing a relative kinematics and dynamics model of a flexible spacecraft containing lumped interference; secondly, designing an expansion interference observer, and estimating the lumped interference in the system model; and finally, introducing the lumped interference estimated value into a system feedback link, and designing an attitude and orbit integrated tracking controller based on the expansion interference observer by combining a backstepping method. According to the method, estimation and compensation of lumped interference can be realized, the spacecraft pose can be tracked to the expected pose of the given target spacecraft by adopting the given control rate, certain dynamic performance of the controller is ensured, and the anti-interference capability and accuracy in the tracking control process of the flexible spacecraft are improved.

The component percentage of the anti-radiation, shielding material is Pb 2-10%, Ba 1-5%, TiO2 5-15%, TiB2 1-10%, the rest is aluminum (Al) Under the smelting temperature, the Al is melt down. Then add the anti-radiation material of Pb, Ba, TiO2 and TiB2, the distribution of the material can be controlled by utilizing the magnetic field and controlling the cooling speed. Thus the grading structure is formed. The material is applied to protect the electrnic component of the spacecraft from outer-sphere radiation damage.

A spacecraft carrier is disclosed. The carrier has a large internal volume for housing at least one spacecraft. The carrier can be used as a repair and maintenance facility in space for spacecraft. Manned and unmanned devices can be stored, repaired and resupplied. The carrier can also transport a number of spacecraft to other locations allowing for an efficient coordinated movement of many spacecraft.

The invention discloses a throwable double-component common-bottom type annular storage tank for a spacecraft. The annular storage tank comprises an annular storage tank body, a center cylinder and aseparation device. A first connecting ring, a second connecting ring and a ring section are symmetrically arranged on the annular storage box body, and the first connecting ring and the second connecting ring are provided with lug pieces. The center cylinder is provided with an upper end frame, a lower end frame and stringers; and the lug pieces and the stringers enable the annular storage tank body and the center cylinder to be connected through bolts to bear external loads together, the upper end frame and the lower end frame are separation surfaces, and the separation device is arranged onthe separation surfaces so that the annular storage tank body can have a throwing function. A connecting ring-single layer common-bottom assembly enables the annular storage tank body to form a double-component structural form, and a fuel agent and an oxidizing agent can be carried at the same time.

The invention relates to a method for detecting random linear network coding tampering data in a spatial information network, relating to the technical field of safety communication of spatial information networks. A source spacecraft node calculates a homomorphic hash function value of a source message, and forwards the source message together with the homomorphic hash function value of the source message in a random linear network coding manner; data packets needing to be subjected to tampering detection are determined self-adaptively; and tampering detection of the data packets needing to be subjected to tampering detection is carried out, so that tampered data packets are found, and abandoned. According to the method disclosed by the invention, tampering detection probabilities of the data packets and the nodes are configured according to different network environments; due to inter-node cooperation, the two tampering detection probabilities are automatically adjusted; the nodes and the data packets needing to be subjected to tampering detection are self-adaptively determined according to the two tampering detection probabilities; therefore, the data packet tampering detection time can be reduced; the transmission distance of a polluted message is shortened; and in addition, specific tampered information in the data packets is found by adopting a binary search method, so that the search time can be obviously reduced.

The invention relates to an impact reducing support structure suitable for a single machine with a small heat conduction amount requirement on a spacecraft, and belongs to the technical field of impact reduction support structures. The impact reducing support structure comprises an upper fastening device, an impact reducing support, a lower fastening device and a connecting embedded part. The impact reducing support provides a threaded mounting hole for the single machine; a single machine mounting foot is a through hole and is connected with the single machine through a fastening screw and apressing gasket; a support interface, connected with the single machine, of the impact reducing support adopts an eccentric design; the connecting embedded part is arranged in a cellular board and provides a threaded mounting hole for the impact reducing support, and the mounting hole of the impact reducing support and the mounting hole of the cellular board are through holes and the impact reducing support and the cellular board are connected with the single machine through fastening screws and pressing gaskets. According to the impact reducing support structure, the step-by-step impact reduction mode is adopted, and the impact reduction efficiency is high; the upper support interface adopts an eccentric design, so that the bearing capacity is high; the impact reducing support is of a cylindrical hollow shell structure, and has the advantage of being low in structural weight and wide in application range.

A method for selecting a magnetic holding relay suitable for improving safety is provided. Aiming at the possible welding failure mode of magnetic holding relay, the magnetic holding relay used in spacecraft is designed to resist surge current shock according to the influence of surge current and surge current duration on relay life at the moment of contact opening and closing. At present, the relay selection and design can only be carried out through the component manual, but only part of the relay type anti-surge current value is given in the component manual, the influence of surge currentand energy generated by surge current duration on relay safety is not clear, and the possibility of over-design or under-design exists in relay selection and design. Compared with the existing designmethod, the method of the invention is more quantitative, more reliable, more operable and has good application value.

The invention relates to an electric transmission rotary joint, and the rotary joint comprises a main shaft, a shell, a motor, a gear ring and N conductive rolling ring sets, wherein N is greater thanor equal to 1. The motor drives the gear ring to rotate, the gear ring is fixedly connected with the shell, and the gear ring drives the shell to rotate, so the shell is a rotating part, the main shaft is located in the shell, the two ends of the main shaft are connected with the shell through the angular contact bearings, the main shaft is connected with a fixing part on the satellite body and is a fixing part, and the N conductive rolling ring sets are located between the shell and the main shaft in parallel and used for achieving an electric transmission structure with an outer rotor and an inner stator fixed. The main shaft fixing and shell rotating mode can achieve series connection of a plurality of rolling ring sets (namely ring body assemblies) and series connection of a pluralityof solar cell arrays, and is particularly suitable for large spacecrafts such as space solar power stations.

The invention relates to an anti-backlash locking mechanism driving control method and system for a spacecraft, and the method comprises the steps: adjusting the driving current of the stepping motor of the anti-backlash locking mechanism in a forward locking mode and a reverse unlocking mode, so that in a forward locking task, the driving torque of the anti-backlash locking mechanism is not larger than the torque during reverse unlocking all the time, locking of the mechanism during operation can be effectively avoided, and the feasibility of the method is verified through test results of special tests.

The invention relates to an asynchronous data synchronization method of a four-machine hot backup real-time system, belongs to the technical field of multi-machine hot backup control systems on spacecrafts, and particularly relates to an asynchronous data synchronization method based on the four-machine hot backup real-time system. Information exchange between single machines is carried out, whether the data information is used or not is determined according to the states of the events received by the four machines, the data information is used as long as the state of the event received by oneof the four machines is valid, the problem that the four machines are asynchronous due to the fact that a single machine cannot receive data caused by faults is solved, and the reliability and safetyof the hot backup system are improved.

The invention relates to the field of auxiliary equipment of spacecrafts, in particular to a reusable stable selfie device and method for an aircraft. The problems that an existing selfie imaging system is low in cost-efficiency ratio and cannot meet the requirement for continuous video monitoring and recording in the execution process of multiple key actions on a satellite are solved. The selfie device comprises a fixing base, a two-axis pointing adjusting module, an ejection assembly, a camera assembly and a control module. A two-axis direction adjusting module and an ejection assembly are adopted to achieve the functions of direction switching and ejection of a camera assembly, and after a selfie task is finished, the camera assembly is reset through a dragging assembly located in the camera assembly to wait for next time of shooting; the single-axis gyrorotor located in the camera assembly is adopted, so that the posture of the camera assembly is basically fixed after ejection, the visual axis of the camera can be ensured to always point to a certain direction or other set fixed airspaces of the aircraft in the separation process of the camera assembly and the aircraft, and continuous images or monitoring videos in the key action execution process are obtained.

The embodiment of the invention provides a spacecraft long-term management stage orbit parameter automatic adjustment method and system. The method comprises an orbit parameter optimization adjustment step of carrying out the optimization adjustment of orbit parameters based on the current target start and stop time of the orbit parameters in a spacecraft long-term management stage, the target orbit periodic change rate at a starting moment, a threshold parameter group, and an extrapolation time amount, generating a corresponding set parameter sequence and an injection orbit parameter of the spacecraft, carrying out difference calculation on the prediction result calculated by the ground control center after the injection orbit parameter is extrapolated by the analogue simulation spacecraft, determining an orbit parameter evaluation sequence used for representing the data characteristics of the multi-dimensional deviation value, and if the evaluation sequence meets the threshold requirement, further optimizing and adjusting the target starting and ending time and the target orbit periodic change rate, and returning to execute the orbit parameter optimization and adjustment step. The method can automatically adjust the orbit parameters of the spacecraft in the long-term management stage, so as to guarantee the stability of the flight attitude of the spacecraft.

The invention discloses a high orbit spacecraft launching method based on multi-rocket combination, and belongs to the technical field of spaceflight launching, and the method comprises the following steps: S1, respectively launching a target spacecraft and a supply spacecraft to a low orbit through rockets; s2, the target spacecraft and the replenishment spacecraft are in butt joint in a low orbit; s3, after docking, the replenishment spacecraft replenishes the propellant carried by the replenishment spacecraft to the target spacecraft, and docking is released after replenishment is finished; and S4, after the target spacecraft obtains propellant supply, the target spacecraft is transferred to the high orbit through power of the target spacecraft. According to the application of the method disclosed by the invention, in a high-orbit large-load spaceflight launching task, a high-thrust carrier rocket with high development and use cost is prevented from being used, and meanwhile, the economic loss caused by failure of a single launching task is also reduced.