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294results about How to "Lower launch costs" patented technology
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Anchor positioning system for detecting planetoid lander
InactiveCN101786504ALower launch costsRisk of drifting away eliminatedCosmonautic vehiclesSystems for re-entry to earthCelestial bodyEngineering
The invention relates to an anchor positioning system for detecting a lander, in particular to the anchor positioning system for detecting a planetoid lander. The anchor positioning system solves the following problems of a telescopic sleeve connecting and positioning system: an anchor bill is inclined when in incidence and transverse impact load is acted on the lander, so as to damage instruments at the inner part of the lander; a moving part of a telescopic sleeve is changed in quality during the incidence, so as to cause difficulty in estimating initial emission energy. An advancing mechanism is fixedly arranged on the side face of a cable box body, a winding mechanism is fixedly arranged on a rear end cover of the cable box body, a locking and unlocking mechanism is fixedly arranged on a right shell body of the winding mechanism, a connecting rope is stored in the cable box body, one end of the connecting rope is fixed on an anchor rod pin, and the other end of the connecting ropeis fixedly connected with a winding reel. A winding motor and an unlocking motor are connected with a pin socket by a lead wire. The anchor positioning system can also be used not only for detecting the lander of a small celestial body with weak surface gravitation such as the comet and the like, but also for automatic operation of a robot in dangerous environments.
Owner:HARBIN INST OF TECH
Single-stage hypersonic vehicle featuring advanced swirl combustion
InactiveUS20080283677A1Reduce complexityShorten the lengthAircraft navigation controlPower plant exhaust arrangementsRamjetLow speed
A single-stage hypersonic vehicle is comprised of a low-speed and a high-speed propulsion system. The low-speed propulsion system propels the single-stage vehicle to a threshold velocity, after which the high-speed propulsion system then takes over. The low-speed propulsion system includes a combined-cycle engine featuring a swirl generator that is integrated into a turbojet engine to provide a compact turbojet and swirl afterburner-ramjet propulsion system. The high-speed propulsion system includes a hypersonic engine that is operable at the threshold takeover velocity and beyond. In various embodiments, the high-speed propulsion system comprises a scramjet, rocket, or scramjet / rocket engine depending requirements. Benefits of the swirl generator design include its ability to rapidly and efficiently atomize, vaporize, mix and burn the fuel and oxidizer for the low speed propulsion system, significantly reduce length, weight, cooling requirements and complexity for both propulsion systems, while maintaining high propulsion performance and reducing propulsion and launch costs.
Owner:UNITED TECH CORP
Novel multi-satellite separating, unlocking and releasing device
InactiveCN103738509ASimple structureEasy to modifyCosmonautic component separationModularityModular design
The invention relates to a novel multi-satellite separating, unlocking and releasing device. Defects of varieties of separating, unlocking and releasing devices in the prior art can be overcome, and the problems that a separating device of a micro spacecraft in multi-satellite carrying and launching process is complex and high in cost are solved. The novel multi-satellite separating, unlocking and releasing device adopts an axial supporting manner and a radial separating manner, good mechanical environment of a carrying micro spacecraft in the launching step is guaranteed, and meanwhile, satellites can be prevented from mutual interference during separation and collisions after separation; the novel multi-satellite separating, unlocking and releasing device is simple in structure, convenient to modify and maintain and high in environmental adaptability, adopts modular design, and has the advantages of low cost, parts simplicity and high reliability.
Owner:BEIHANG UNIV
Folding lightweight landing mechanism
InactiveCN102092484ASimple configurationSimple designSystems for re-entry to earthCosmonautic landing devicesInstrumentationAsteroid
Owner:HARBIN INST OF TECH
Modular spatial curved surface folding and unfolding antenna mechanism based on rib mechanisms
ActiveCN104022337ARealize modular expansionRealize modular productionCollapsable antennas meansComputer moduleSatellite
The invention relates to an antenna mechanism, in particular to a modular spatial curved surface folding and unfolding antenna mechanism based on rib mechanisms. The modular spatial curved surface folding and unfolding antenna mechanism aims to solve the problems that an existing unfolding-type parabolic cylinder antenna structure is low in molded surface precision, poor in stability and low in integral rigidity, an umbrella-shaped antenna structure is not suitable for reflecting surface stretch forming, and a truss antenna structure capable of being unfolded only can be unfolded unidirectionally but a parabolic cylinder can not be formed. The modular spatial curved surface folding and unfolding antenna mechanism comprises two speed-controlled motors, two longitudinal connection rib mechanisms, two folding and unfolding connection mechanisms, a plurality of transverse connection rib mechanisms, a plurality of single rib mechanisms, a plurality of single modular folding and unfolding mechanisms, a plurality of tensioning pull ropes and a plurality of modular connection mechanisms. Two transverse outer side supporting frames are arranged between the two longitudinal connection rib mechanisms in parallel and are respectively formed by connecting multiple transverse connection ribs in series, and the two folding and unfolding connection mechanisms, the single modular folding and unfolding mechanisms and the single rib mechanisms are arranged in the supporting frames. The modular spatial curved surface folding and unfolding antenna mechanism is applied to satellites, space stations and space probes.
Owner:HARBIN INST OF TECH
Integrated shock absorbing vibrating isolation device for stellite
InactiveCN1911734ASo as not to damageLower launch costsCosmonautic vehiclesCosmonautic partsStellite alloyIsolating mechanisms
Owner:郑钢铁 +3
Magnetic liquid damp shock absorber
The invention relates to a magnetic liquid damp shock absorber and belongs to the field of mechanical engineering vibration. The magnetic liquid damp shock absorber solves the problems that under the condition that external amplitude is larger than a gap between a shell body and a permanent magnet of an existing magnetic liquid damp shock absorber, the permanent magnet collides with an inner wall of the shell body and thus service life of the shock absorber is shortened and shock absorption performance is influenced. The permanent magnet is large in specific gravity and difficult to process to a required shape. A cylindrical permanent magnet of the magnetic liquid damp shock absorber is a rubber cylindrical permanent magnet (7). A wedge angle gasket (5) is arranged at the bottom inside the shell body (4) and a V-shaped included angle of the wedge angle gasket (5) faces upwards. The rubber cylindrical permanent magnet (7) absorbing magnetic liquid (6) is arranged in the shell body (4) and is magnetized along the axial direction. The axis of the rubber cylindrical permanent magnet (7) is perpendicular to a horizontal plane. The V-shaped included angle of the wedge angle gasket (5) is 60 degrees to 170 degrees. The gap between the inner wall of the shell body (4) and the rubber cylindrical permanent magnet (7) is smaller than or equal to the amplitude of shock absorption required by the shock absorber. The magnetic liquid damp shock absorber is simple, reliable, obvious in shock absorption effect and long in service life.
Owner:BEIJING JIAOTONG UNIV
Parameter selection method adopting onboard control moment gyroscope group vibration-isolating platform
InactiveCN102778891AOvercome wasteAchieve vibration isolationAttitude controlSpacecraft attitude controlEngineering
The invention relates to a parameter selection method adopting an onboard control moment gyroscope group vibration-isolating platform and belongs to the field of the attitude control and the vibration and shaking control of a spacecraft. (1) an idea of the cube vibration-isolating platform is adopted, the cube vibration-isolating platform is arranged between a configuration consisting of a plurality of control moment gyroscopes and a satellite body and the high frequency vibration caused by a plurality of control moment gyroscopes is isolated to the greatest extent; and (2) by reasonably selecting parameters of the onboard control moment gyroscope group vibration-isolating platform, the problem of effectiveness loss of other onboard systems, which is caused after the vibration-isolating platform is applied on a satellite, is avoided.
Owner:BEIJING INSTITUTE OF TECHNOLOGYGY
Curve shaped bullet holder in autosegregation
InactiveCN1687693AImprove emission efficiencyLower launch costsAmmunition projectilesCartridge ammunitionCurve shapeEngineering
The invention is the curved self-separating bullet pallet, and relates to the improvement of the pallet structure. The invention aims to solve the unstable separation of the straight bullet pallet. The sections of the left distinguish 1 and the right distinguish 2 are two corresponding semicircles; the upper part of one side of the inner face 11 of the left distinguish 1 opens the left bullet hole 3; the inner face 11 of the left distinguish 1 has left tooth profile 5; the upper part of one side of the inner face 12 of the right distinguish 2 opens the right bullet hole 4 that is corresponding to the left bullet hoe 3; the inner face 12 of the right distinguish 2 has the right tooth profile 6 that engages with the left tooth profile 5; the left distinguish 1 is the curved body 9 and the right distinguish 2 is the curved body 10 corresponding to the left distinguish. The invention improves the launching efficiency and lowers the cost.
Owner:HARBIN INST OF TECH
Fuel-free spacecraft propelling system based on spatial atomic oxygen and propelling method
InactiveCN102767497ARealize on-orbit full life cycle workAchieve normal workCosmonautic vehiclesCosmonautic partsSpacecraft propulsionElectricity
Owner:BEIJING INST OF SPACECRAFT ENVIRONMENT ENG
Fully compliant pneumatic mechanical arm structure
InactiveCN106584507ASimple structural designSimplify the assembly processToolsArmsHardnessFiber-reinforced composite
The invention discloses a fully compliant pneumatic mechanical arm structure which adopts a lengthwise conical structure. The fully compliant pneumatic mechanical arm structure comprises a fully compliant pneumatic mechanical arm main body, a mechanical arm center body positioned in the conical center hole of the fully compliant pneumatic mechanical arm main body, and air passages evenly surrounding the mechanical arm center body and in axially symmetrical distribution, wherein the air passages extend from a large-diameter end to a small-diameter end and keeps a certain distance with an end surface so as to keep gas inside the air passages, and a layer of fiber reinforced composite material is arranged on the outer surface of a mechanical arm; the air passages in axially symmetrical distribution are of lengthwise conical structures; the fully compliant pneumatic mechanical arm main body is made of a super elastic material, more than 200 percent of deformation is produced under pneumatic pressure, and the hardness of the mechanical arm center body is higher than that of the fully compliant pneumatic mechanical arm main body. Compared with the prior art, the fully compliant pneumatic mechanical arm structure provided by the invention imitates octopus tentacles, and integrates the air passages and the mechanical arm, so that the structural design, the processing and manufacturing, and the assembling of the entire mechanical arm are greatly simplified, 3D printing technology is used for one step processing, and objects in complex shapes can be grabbed, the mass is light, and the transmitting cost is reduced.
Owner:BEIJING INST OF SPACECRAFT ENVIRONMENT ENG
Separation modular satellite system and method based on medium earth obit (MEO) data relay
ActiveCN102932050ARealize real-time data transmissionNo significant increase in costRadio transmissionEnd-to-end delaySatellite
The invention discloses a separation modular satellite system and a method based on MEO data relay. The system comprises a separation modular satellite cluster, an MEO data relay system and an earth station, wherein the separation modular satellite cluster generates data after onboard processing and sends the data to the MEO data relay system; the MEO data relay system performs routing on the data and sends the data to the earth station in a transparent forwarding mode; and the earth station performs further analysis processing on the received data to obtain a final result. According to the system and the method, end-to-end delay of the separation modular satellite system can be effectively reduced, and therefore real-time data transmission of the separation modular satellite system is achieved.
Owner:BEIJING UNIV OF POSTS & TELECOMM
Power control unit (PCU) control system based on bidirectional multi-port converter with wide voltage range
ActiveCN105391301AReduce volumeReduce weightDc-dc conversionPhotovoltaic energy generationMicrochiropteraElectrical battery
The invention provides a spacecraft power controller control system based on a bidirectional multi-port converter with a wide voltage range. The spacecraft power controller control system comprises a main power part, a field programmable gate array (FPGA) part, a digital-to-analog conversion part, a voting circuit and a controlled volume sampling circuit, wherein the main power part is composed of N main power conversion modules, each main power module adopts a bidirectional three-port converter to achieve adjustment and scheduling of energy among a solar array SA, a bus BUS and a battery BAT, and the N main power modules are connected in parallel to achieve power expansion; the FPGA part comprises N FPGA modules corresponding to the N main power modules; the controlled volume sampling circuit is used for sampling the controlled volume of the power part and providing the controlled volume to the FPGA part through the A / D conversion part; driving signals given out by the FPGA are used for respectively controlling the N main power modules through the D / A conversion circuit; and the FPGA part serves as a control core of the system, and the control function of a PCU is achieved by the converters of the main power modules through a control algorithm and a control logic.
Owner:SHENZHEN AEROSPACE NEW POWER TECH
Spacecraft solar panel conformal antenna
ActiveCN106848558AWide range of bandwidthGuarantee normal optical remote sensing functionAntenna supports/mountingsRadiating elements structural formsInterstellar probeConformal antenna
On the basis of defining the concept of a spacecraft solar panel conformal antenna, the invention discloses S1, a spacecraft solar panel conformal antenna structure; S2, a spacecraft solar panel conformal antenna array structure; S3, a spacecraft solar panel conformal antenna array feeding mode; S4, a spacecraft solar panel conformal antenna beam control network; S5, a spacecraft solar panel conformal antenna microwave signal transmission line; S6, a connecting method of the spacecraft solar panel conformal antenna and a spacecraft. The spacecraft solar panel conformal antenna has the advantages of reducing the spacecraft launching cost, prolonging the in-orbit service life of the spacecraft, reducing the fairing requirement, facilitating appearance attractiveness of the spacecraft, improving spacecraft maneuverability and safety, and being high in replicability and good in expandability. Multiple solar panel conformal antennas with different orbit heights and types can be applied through networking, and cover with a bigger bandwidth range and a wider spatial range can be achieved through mission planning. Besides, the conformal antenna thought can be expanded to the field of interstellar probe antennas. The spacecraft solar panel conformal antenna has a wide application prospect.
Owner:耿文东
Wireless test launch and control system for aircraft
ActiveCN110412910AOptimize layoutEnsure safetyProgramme controlComputer controlTelecommunications linkCommand and control
The invention discloses a wireless test launch and control system for an aircraft. The wireless test launch and control system is composed of ground equipment and rocket-borne equipment, wherein the ground equipment comprises a command and control machine, a ground data link, a telemetering station, a switch and a display machine, and the rocket-borne equipment contains a rocket-borne data link, arocket-borne telemetering instrument and a rocket-borne computer. A wireless communication link contains an upward data link access, a downward data link access and a downward telemetering access andrealizes a test launch and control instruction, test launch and control feedback and telemetering data transmission of the aircraft. The rocket-borne data link has a low-power-consumption sleep function, can wake up a rocket-borne battery of the aircraft and control the rocket-borne battery of the aircraft to be activated as well as power distribution after an upward wireless test launch and control instruction is received and then starts a test launch and control process. The system disclosed by the invention does not need to connect ground test launch and control equipment with the aircraftby virtue of an umbilical cable, so that ablation of the umbilical cable due to flame of an engine of the aircraft is avoided; and ground communication equipment is simple to pave, a target range communication network does not need to be used, and the system disclosed by the invention has a function of receiving telemetering data in a flying process of the aircraft and a command and control function at the same time.
Owner:CHINA ACAD OF AEROSPACE AERODYNAMICS +1
Solar saucer-shaped floating aircraft provided with shrinkable airbag
InactiveCN105947169AReduce volumeSolve the problem of susceptibility to airflow interferencePower plant typeGas-bag arrangementsPropellerAerospace engineering
The invention discloses a solar saucer-shaped floating aircraft provided with a shrinkable airbag. The solar saucer-shaped floating aircraft provided with the shrinkable airbag comprises a saucer-shaped aircraft body which is transformed from a saucer-shaped motor, and a propeller is arranged in the middle of a rotor of the saucer-shaped aircraft body. The shrinkage airbag is installed on the edge of the saucer-shaped aircraft body, and a solar power generation device is installed on the skin of the upper portion of the airbag. The lifting force of the aircraft body can be provided by air in the airbag and the propeller in the middle of the aircraft body independently or jointly, and power, in the horizontal direction, of the aircraft body is provided by a steering engine. The shrinkage airbag is mainly composed of airbag skin, a pull pressing rod system in the airbag and a telescopic system supporting the skin. The solar power generation device is in a fan shape, is circularly distributed on the skin in an attached mode, and moves along with the stretching or retracting of the skin.
Owner:王庆方 +1
Short-range laser defensive system
PendingCN107514936ALow strike abilityLower launch costsDirected energy weaponsUncrewed vehicleEngineering
The invention discloses a short-range laser defensive system. The short-range laser defensive system comprises an integrated complete machine cabinet, a fiber laser arranged inside the integrated complete machine cabinet, a controller arranged inside the integrated complete machine cabinet, a search radar arranged on the end face of the top of the integrated complete machine cabinet, a follow-up rotating table arranged on the end face of the top of the integrated complete machine cabinet, a laser emission barrel arranged on the follow-up rotating table and a photoelectric tracker arranged on the laser emission barrel; and the laser output end of the fiber laser is connected with the laser input end of the laser emission barrel, and the search radar, the follow-up rotating table, the fiber laser, the laser emission barrel and the photoelectric tracker are all connected with the controller in an electric control manner. According to the short-range laser defensive system, the striking capacity on low, slow and small unmanned aerial vehicles, the very high short-range low-altitude defense capability is very high, the structure is compact, and transportation is convenient.
Owner:合肥正阳光电科技有限责任公司
A kind of preparation method of fiber reinforced ceramic matrix composite material with low thermal expansion coefficient
ActiveCN106507851BRealize controllable adjustmentSimple preparation processChemical vapor deposition coatingCarbon fibersThermal expansion
The invention discloses a method for preparing a fiber-reinforced ceramic matrix composite material with a low thermal expansion coefficient, comprising: firstly, carrying out structural weaving of a fiber preform, and the material of the fiber preform is carbon fiber or silicon carbide fiber, or a combination of the two ; secondly, the interfacial phase is prepared on the surface of the fiber preform; thirdly, the chemical vapor deposition process is used first, and then the fiber preform is densified by the impregnation and cracking process of the organic precursor. The invention can obtain the fiber-reinforced ceramic matrix composite material with low coefficient of thermal expansion.
Owner:SHANGHAI INST OF CERAMIC CHEM & TECH CHINESE ACAD OF SCI
Yawing maneuvering control method based on sinusoidal yawing guidance principle
ActiveCN102880059AGuaranteed supplyReduce capacity requirementsAttitude controlAdaptive controlMomentumControl system
The invention discloses a yawing maneuvering control method based on a sinusoidal yawing guidance principle. The sinusoidal yawing guidance principle is designed in the yaw axis maneuvering method of an aircraft, thus, a yaw axis follows a sinusoid and the rotating speed of a sailboard is calculated based on the sinusoid at the same time, so as to ensure that an inclined angle between the sun direction and the normal of the sailboard is minimum in this mode. With the adoption of the sinusoidal yawing guidance principle, the energy of the aircraft is ensured, the demand of original yawing maneuvering mode to the capacity of an angular momentum exchange device is reduced, the demand on the capacity of the angular momentum of the aircraft is reduced, and a method is provided for reducing the weight of a control system.
Owner:BEIJING INST OF CONTROL ENG
Support-integrated reflector
The invention relates to a structural component of a space camera, and in particular to a support-integrated reflector. The support-integrated reflector comprises a reflector panel integrally formed by a metal material, a support plate and a back plate. One side of the support plate is the reflector panel, and the other side of the support plate is the back plate. The support plate comprises an outer frame, an inner frame and net-shaped ribs, the outer frame is positioned on the outer circumference of the support plate, the inner frame is positioned at the center of the support plate, and the net-shaped ribs are positioned between the outer frame and the inner frame. The outer frame is provided with a plurality of sub-frame installing plates and a plurality of flexible brackets. The support-integrated reflector is capable of solving the technical problem that the traditional reflector and the support structure thereof need to be machined, assembled and used alone. The reflector and the support are integrally designed, the traditional reflector support structure is eliminated, and one component is used as the reflector and the support. The weight of the whole device and the number of the parts can be reduced, the reflecting cost is decreased, and the rigidity of the integrated structure is improved.
Owner:XI'AN INST OF OPTICS & FINE MECHANICS - CHINESE ACAD OF SCI
Collapsing and unfolding support tubes for space membrane facilities
InactiveCN105474780BImprove mechanical propertiesGood dimensional stabilityCosmonautic vehiclesLayered productsResin-Based CompositeLow density
The invention relates to aerospace vehicle equipment, and discloses a retractable and unfolded support tube for a space film facility, comprising: an upper thin-walled shell [1] and a lower thin-walled shell [2] with a cross-section in the shape of "Ω". The semicircular concave surfaces are opposite to each other, and the flat edges on both sides are glued to each other to form a pod-shaped thin-walled support tube. During the launch stage of the spacecraft, the thin-walled support tube is flattened into a flat strip and curled around itself or a drum. After the spacecraft is launched to the predetermined orbit, the thin-walled support tube unfolds and returns from the flat roll state to the long tube state. The above-mentioned upper thin-walled shell [1] and lower thin-walled shell [2] are prepared from high specific modulus fiber-reinforced resin matrix composite materials. The invention solves the problems of large volume, complex structure, low deployment stability and reliability of the original structure, and has the advantages of large elastic recovery force, high elastic recovery rate, good elastic recovery reliability, low density, high strength, high rigidity and low cost. inferior advantages.
Owner:SHANGHAI AEROSPACE SYST ENG INST
Composite thermosetting film for inflated spatial expanded structure and its making and rigidizing process
The present invention relates to composite film and its making and rigidizing process. The composite thermoset film for inflated spatial expanded structure includes inner gas blocking layer, rigidized layer, heating layer, and outer heat insulating layer successively from inside to outside. Fiber fabric woven in two or three directions is soaked in thermoset resin to form the pre-soaked fabric; and metal foil or metal filament is set between two layers of polyimide film with thermoset adhesive painted in the inner layer and formed via heating and pressurizing in a hot press. Perforated polyimide film or polyester film with aluminum plated to the one or two sides is adhered to dacron net and formed into shape the same as the required inner gas blocking layer. The composite thermoset film of the present invention has the advantages of light weight, small occupied emitting volume, lowered emitting cost and high reliability.
Owner:HARBIN INST OF TECH
Satellite power controller
ActiveCN106647920AImprove efficiencyImprove reliabilityPhotovoltaic energy generationElectric variable regulationCapacitancePower controller
The invention provides a satellite power controller. The controller is used for tracking the peak power of N photovoltaic cell arrays which are in parallel connection. The controller comprises N output diodes, a first switch S1, a second switch S2, a third switch S3, a second diode D2, a third diode D3, an output capacitor C1, an output capacitor Co, an inductor L1, an input voltage sensor, an input current sensor, an output voltage sensor and a control unit. The satellite power controller combines the S3R sequential switch shunting regulation and MPPT maximum power point tracking control for enhancing the satellite power system peak power tracking efficiency and reliability, thus obtaining a satellite platform with high power, high efficiency and high reliability.
Owner:SHANDONG INST OF AEROSPACE ELECTRONICS TECH
Power density adjustable film reflecting and condensing space solar energy collecting station
InactiveCN103868246ASolve the weight problemSolve the costSolar heating energySolar heat devicesOptical axisPlane mirror
The invention discloses a power density adjustable film reflecting and condensing space solar energy collecting station, which comprises a spacecraft platform, an optical system and a light path adjusting mechanism, wherein the optical system comprises a main condenser, an auxiliary condenser and a plane mirror; the main condenser and the auxiliary condenser are arranged in opposite directions, optical axes of the main condenser and the auxiliary condenser are overlapped, and a normal distance between the main condenser and the auxiliary condenser is adjusted through the light path adjusting mechanism to change the power density of output light beams; the angle of the plane mirror is adjusted through the light path adjusting mechanism to change the direction of the output light beams; sunlight is condensed and reflected to the auxiliary condenser through the main condenser, the reflected light rays are condensed and reflected to the plane mirror through the auxiliary condenser, and the light beams are reflected to a target through the plane mirror.
Owner:CHINA ACADEMY OF SPACE TECHNOLOGY
Efficient isotope battery based on gas radioactive source
ActiveCN105788692AReduce the effective areaIncrease light intensityRadiation electrical energyElectrical batteryIsotope
The invention discloses an efficient isotope battery based on a gas radioactive source. The isotope battery comprises a battery base with an embedment groove, a radioactive light source support put in the embedment groove of the battery base, at least one radioactive source radiation light-emitting unit put on the radioactive light source support, a lens buckled on the embedment groove of the battery base and isolating the inside and the outside of the embedment groove, an output end support connected with the battery base, and a photovoltaic conversion unit put on the output end support, matched with the lens in position and used for outputting electric energy, wherein light rays emitted by all the radioactive source radiation light-emitting units are projected to the photovoltaic conversion unit after being gathered by the lens. A gas isotope radioactive source is ingeniously used for stimulating a radioactive light-emitting material to emit light, light rays are gathered on the photovoltaic conversion unit by the lens through the combination of an optics principle so that electric energy can be output, the light intensity on the photovoltaic conversion unit is greatly improved, the efficiency of the photovoltaic conversion unit is greatly improved, and the output power and the energy conversion efficiency of the isotope battery are improved.
Owner:MATERIAL INST OF CHINA ACADEMY OF ENG PHYSICS
Connecting and separating device for satellite batch launching
ActiveCN111086658ASafe and reliable launchOn-orbit unlocking is safe and convenientCosmonautic vehiclesCosmonautic component separationStructural engineeringRocket
The invention discloses a connecting and separating device for satellite batch launching. The connecting and separating device comprises a base, docking mechanisms and a locking mechanism. The dockingmechanisms are connected with the satellite body, the multiple docking mechanisms are arranged in an overlapped mode, elastic mechanisms are arranged between two adjacent docking mechanisms and between the bottommost docking mechanism and the base, the two adjacent docking mechanisms are connected in an abutting mode through the corresponding elastic mechanism, and the bottommost docking mechanism is connected with the base in an abutting mode through the corresponding elastic mechanism. The base is connected with the docking mechanism through the locking mechanism, and the locking mechanismis used for fixing the docking mechanism to the base. By adopting the scheme provided by the invention, large-batch satellite launching can be simultaneously realized, the launching cost is greatly reduced, and guarantee is provided for satellite large-batch launching and constellation construction; the design meets the expandability of one-rocket multi-satellite launching and the uniformity of satellite design and manufacturing, all satellites are convenient to assemble and loose in requirement before launching, the safety requirement is met in the launching process, and on-orbit unlocking issafe and convenient.
Owner:中国星网网络应用有限公司
Spacecraft attitude control method for controlling moment gyro singularity avoidance
ActiveCN105223961APrecise torque outputLower launch costsAttitude controlPosition/course control in three dimensionsSingularity avoidanceLinear model
The invention discloses a spacecraft attitude control method for controlling moment gyro singularity avoidance, specifically comprising the steps as follows: S1, solving an SGCMG (single gimbal control moment gyro) group angular momentum set; S2, deciding the angular momentum amplitude of an SGCMG group and spacecraft system; S3, determining the optimal control performance index of a spacecraft attitude control system; S4, determining a linear model of the SGCMG group and spacecraft system; S5, determining the linearization range of the SGCMG group and spacecraft system; S6, determining the linearization range meeting the constraints; and S7, using nonlinear predictive control to implement spacecraft attitude control. By using the method, the attitude of a spacecraft can be controlled precisely, the spacecraft launching cost is reduced, and the in-orbit operation life of spacecrafts is increased.
Owner:BEIJING MECHANICAL EQUIP INST
Fuel-Free Spacecraft Propelling System Based on Spatial Atomic Oxygen and Propelling Method
InactiveUS20150210406A1Lower launch costsCosmonautic vehiclesCosmonautic propulsion system apparatusSpacecraft propulsionOxygen plasma
A fuel-free spacecraft propelling system having an open-ended outer cylinder of a propelling device and an atomic oxygen collecting device is disclosed. The latter is arranged at the forwardly-propelled front end of the outer cylinder and is hermetically connected with an RF generating device and an ion cyclotron wave heating device through a magnetic confinement device. A spiral wave discharge oxygen plasma inlet and a spiral wave discharge oxygen plasma outlet in the ion cyclotron wave heating device are respectively provided with another magnetic confinement device. The propulsion of the invention does not need to carry the propellant, which greatly reduces the launch costs, and enables a spacecraft to advantageously have an increased orbit life over existing spacecraft systems.
Owner:BEIJING INST OF SPACECRAFT ENVIRONMENT ENG
Light mechanical collision type space debris capturing web claw
The invention discloses a light mechanical collision type space debris capturing web claw. The light mechanical collision type space debris capturing web claw comprises a bottom plate and a pluralityof web claw teeth, wherein the plurality of web claw teeth are mounted on the bottom plate; toothed claw switch module for driving the web claw teeth to be closed when the web claw teeth have collisions with a space debris is also mounted on the bottom plate; the web claw teeth consist of claw teeth with spokes on the two sides; and when the web claw teeth are closed, the plurality of web claw teeth and the spokes on the two sides of the plurality of web claw teeth form a semi-closed capturing envelope to prevent a capturing target from escaping. According to the light mechanical collision type space debris capturing web claw, at the moment of approaching the space debris in the capturing process, the capturing envelope is formed by triggering the switch through triggering a collision triggering switch of the bottom plate and releasing a torsion spring to drive the closed claw teeth; and therefore, the motion of a non-cooperative target is restricted and limited to finally realize thecapturing.
Owner:RES & DEV INST OF NORTHWESTERN POLYTECHNICAL UNIV IN SHENZHEN +1
Clamping and releasing device for on-orbit mutual separation of master spacecraft and slave spacecraft
ActiveCN106628273AHigh clamp release reliabilityHigh separation precisionCosmonautic vehiclesCosmonautic component separationEngineeringSlide plate
The invention provides a clamping and releasing device for on-orbit mutual separation of a master spacecraft and a slave spacecraft. The clamping and releasing device comprises a pin puller (7), a pin puller support (12), two springs (8), a pre-tightening bolt support (9), a pre-tightening bolt (10), a thrust spring (4), four polytetrafluoroethylene sliding blocks (32), two door type sliding plate limiting supports (6), a door type sliding plate (2), two tension spring supports (11), a driven separation plate (18), a driving push plate (17), a main installation plate (1) and a thrust spring installation support frame (5). The middle of the main installation plate (1) is provided with a square hole (29). The thrust spring installation support frame (5) is fixed to the square hole (29). One end of the thrust spring (4) abuts against the thrust spring installation support frame (5). The door type sliding plate (2) is clamped in a sliding groove (33). The pin puller (7) is installed on the pin puller support (12). The driving push plate (17) is wrapped in the driven separation plate (18).
Owner:NAT SPACE SCI CENT CAS
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