36results about How to "Lower launch costs" patented technology

The invention relates to a non-pyrotechnic secondary satellite and rocket locking and separating device which comprises a shape memory alloy primary separating assembly, a pressing and releasing assembly and a turntable four-claw secondary clamping and separating assembly, and the shape memory alloy primary separating assembly is arranged in a box body of the pressing and releasing assembly; one end, extending out of the shell, of a grooving bolt shaft of the shape memory alloy first-stage separation assembly penetrates through a through hole of the separation block to be connected with a nut, and the shape memory alloy rod provides driving force for the shape memory alloy first-stage separation assembly to jointly form a first-stage unlocking mechanism; a clamping jaw guide groove of the rotary table four-jaw two-stage clamping and separating assembly is fixed to the bottom plate, a rotary table controls a clamping jaw set to be locked, and clamping jaws are connected with a separating block of the pressing and releasing assembly to jointly form a two-stage unlocking mechanism. When the device is locked, after being locked by the shape memory alloy first-stage separation assembly, the device is locked by the turntable four-claw second-stage clamping separation assembly; when the device is separated, the device is unlocked by the turntable four-claw secondary clamping and separating assembly, and then is unlocked by the shape memory alloy primary separating assembly.

The invention discloses a fine tracking detection light path structure for a free space laser communication terminal, and belongs to the technical field of satellite laser communication. The system comprises a receiving light path convergence lens, an inclined end face optical fiber, a relay amplification lens and a fine tracking detector. A fiber core of the inclined end face optical fiber is located at the focus of the receiving light path convergence lens, and the fiber core of the inclined end face optical fiber and the center point of the fine tracking detector meet the object image conjugate relation with the relay amplification lens. The optical axis of the receiving light path convergence lens, the optical axis of the relay amplification lens, the optical axis of the inclined end face optical fiber and the normal of the inclined end face optical fiber are in the same plane; reflected light beams of the optical fiber with the inclined end face are collected through the relay amplification lens and imaged to the high-resolution area array detector, high-precision detection on the angle of the received laser beams is achieved under the condition that extra laser energy loss is not introduced, and the result is not affected by angle drifting of all optical elements of a receiving branch.

The invention discloses a maintainability modeling, distribution and prediction method based on reuse. The method comprises the following steps: establishing a hierarchical structure of a reuse carrier rocket; constructing a judgment matrix of the scheme layer; weight vectors of the criterion layer and the scheme layer are calculated, and consistency check of the judgment matrix is carried out; calculating a combined weight vector of each scheme in the scheme layer according to the weight vectors of the criterion layer and the scheme layer; obtaining the maintenance degree of each scheme in the scheme layer according to the combined weight vector; determining maintainability prediction parameters and a reference unit; calculating the average repair time of the subsystems; calculating the average restoration time interval of the whole rocket and the uncertainty of the average restoration time of the whole rocket; and predicting a maintenance degree index of the target layer. The method solves the problems of lack of distribution data, difficult prediction process, difficult prediction result quantification and uncertainty and the like of the reusing carrier rocket maintainability index under the background of minimum subsamples.

The invention relates to a modularized pressing and releasing device and method applied to stacked satellites. The modularized pressing and releasing device comprises a force bearing base, an off-axis unfolding and locking device, a tension spring driving device, a pressing cover assembly and a hot work pressing and unlocking assembly. The initiating explosive pressing and unlocking assemblies are installed on the left side and the right side of the upper face of the force bearing base correspondingly, the off-axis unfolding and locking devices are installed on the portions, located right in front of the two initiating explosive pressing and unlocking assemblies, of the force bearing base correspondingly, and the off-axis unfolding and locking devices are connected with the initiating explosive pressing and unlocking assemblies; the upper end of the fire work pressing unlocking assembly is connected with the pressing cover assembly, the two edge openings of the front end of the force bearing base are both connected with the tension spring driving devices, and the other ends of the tension spring driving devices are connected with the pressing cover assembly. The modularized pressing and releasing device applied to the stacked satellites has the advantages that batch production is facilitated, modularized design is achieved, and good interchangeability is achieved.

The utility model discloses a pneumatic ejection launching device which comprises a launching cylinder and a piston arranged in the launching cylinder. The launching barrel comprises a barrel body and a buffering section arranged at the front end of the barrel body, a weakening structure is arranged on the buffering section in the circumferential direction of the buffering section, a blocking structure used for limiting movement of the piston is arranged in the buffering section, and when the piston moves to the position of the blocking structure and is blocked, the piston is blocked. And the buffer section is damaged at the position of the weakening structure under the action of the piston. The buffer section is arranged at the front end of the launch canister body to limit the piston to move along with the aircraft, so that the aircraft can be quickly separated from the piston after being launched out of the launch canister, and the influence on flight of the aircraft is reduced.

This invention discloses a space radiating plate based on a large vapor chamber and its manufacturing method, belonging to the field of spacecraft thermal control technology. The radiating plate includes a one-piece sealed metal cavity, the interior of which is a vacuum vapor chamber. The cavity has opposing first and second sidewalls; the outer surface of the first sidewall forms a radiating surface, and the inner side of the second sidewall integrates cooling channels. A micro-pillar array structure, directly formed integrally through micromachining, is disposed on the inner wall surface of the vacuum vapor chamber, and the cavity is filled with a phase change working fluid. This invention highly integrates the vapor chamber, channels, and radiating surface, utilizing phase change heat transfer to achieve ultra-high thermal conductivity and temperature uniformity, solving the problems of low heat dissipation efficiency, anisotropy, large weight, and poor adaptability to the space environment in existing technologies. It is suitable for thermal management of high-power spacecraft such as satellites, space stations, and space data centers.

The utility model discloses an inclined nozzle angle engine shell suitable for a high-altitude aircraft, which comprises a shell assembly, a top cover assembly is matched with the top of the shell assembly, a bottom sealing cover is matched with the bottom of the shell assembly, four inclined nozzles are arranged on one side of the bottom sealing cover, and the four inclined nozzles are uniformly distributed. According to the engine shell with the inclined spray pipe angle suitable for the high-altitude aircraft, the spray pipe is obliquely arranged, the direction of hot airflow sprayed by the spray head is changed, the hot airflow sprayed downwards originally is sprayed towards the side face, and therefore the hot airflow sprayed towards the side face is changed into the hot airflow sprayed towards the side face. Damage of sprayed hot air flow to the launching pad and surrounding objects can be greatly reduced, launching cost is reduced, meanwhile, the reflecting plate is arranged below the spraying pipe, most of heat emitted by the spraying pipe can be reflected to the outer side perpendicular to the spraying pipe, and the influence of the heat of the spraying pipe on the surrounding objects is further reduced.

The invention discloses a space plasma energy collection device and method based on spacecraft surface potential difference, and belongs to the technical field of spacecraft energy supply. According to the invention, each sensor probe is connected with each monitoring point and is used for monitoring the potential of the monitoring point; the input end of the processor is connected with the output end of the sensor probe; the output end of the processor is connected with the first input end of the voltage regulator; when the potential of a certain monitoring point reaches a discharge threshold value, the monitoring point is conducted to work, and the storage battery is charged by utilizing the potential difference between the potential of the monitoring point and the potential before the grounding reference point; the monitoring point and the grounding reference point are connected with the input end of the phase inverter; and the output end of the phase inverter is connected with the second input end of the voltage regulator and the storage battery. Accumulation of charges on the surface of a spacecraft can be reduced, the discharge frequency of the plasma contactor is reduced, and the structure is compact.

The invention discloses a satellite propulsion device and a control method thereof, the satellite propulsion device comprises a shell, a drainage core body, an oscillation atomization assembly, a heating assembly, a humidity monitoring assembly and an electric propulsion controller, and the shell is hollow to form a storage chamber for storing a propellant; a guide hole channel is formed in the side wall of the material storage chamber; the drainage core body is arranged in the storage chamber; the end part of the drainage core body is inserted into the guide hole channel; the oscillation atomization assembly is connected with the shell, is arranged at a port of the guide hole channel, is in contact with the drainage core body and is used for atomizing and spraying out the propellant on the end face of the drainage core body; the heating assembly is connected with the shell, is arranged on the outer wall of the shell and is used for melting the propellant; the humidity monitoring assembly is connected with the shell and is arranged on a path of the liquid particles sprayed by the oscillation atomization assembly; the electric push controller is arranged on the shell; the electric push controller is electrically connected with the oscillation atomization assembly, the heating assembly and the humidity monitoring assembly. Attitude adjustment is achieved by heating and atomizing the propellant, and the satellite operation cost is reduced.

The present application relates to the field of spacecraft thermal control technology, and provides a two-phase fluid loop system for spacecraft heat dissipation, comprising an evaporation unit, a pump-driven two-phase fluid loop and a heat pump loop, the evaporation unit is provided with a liquid accumulator for storing heat transfer working medium, a micro-channel cold plate, the micro-channel cold plate comprises a micro-channel cold plate inlet, a micro-channel cold plate liquid phase outlet and a micro-channel cold plate vapor phase outlet; the pump-driven two-phase fluid loop is provided with a mechanical pump and a first radiator, the outlet of the first radiator is connected with the liquid accumulator liquid phase fluid inlet of the liquid accumulator; the heat pump loop is provided with a compressor and a second radiator, and the second radiator is connected with the liquid accumulator through a throttle valve. The present application can effectively reduce the weight of the system and the launch cost, and the micro-channel cold plate has the function of vapor-liquid separation in a closed system, which reduces the internal flow instability, improves the heat exchange capacity and the critical heat flux density, and solves the problems of mechanical pump cavitation and compressor liquid knock.

The utility model discloses a dual-wavelength optical path structure capable of reducing ASE noise, which belongs to the field of aerospace optical amplifiers and is used for reducing the ASE of light output by an optical amplifier, improving the OSNR and improving the sensitivity of a system. The device mainly comprises a jumper IN, a secondary amplification light path and an optical fiber ER, the jumper IN is connected with a wavelength division multiplexer M1, the wavelength division multiplexer M1 is connected with the input end of a wavelength division multiplexer W2 and the input end of a wavelength division multiplexer W5, the wavelength division multiplexer M1 screens two kinds of light with different wavelengths to form two light paths, and the two light paths are connected with the optical fiber ER. The output end of the wavelength division multiplexer W2 and the output end of the wavelength division multiplexer W5 are sequentially connected with the optical fiber ER and the filters BP, the output ends of the two filters BP are connected with the wavelength division multiplexer M2, the two optical paths are converged on the wavelength division multiplexer M2, and the wavelength division multiplexer M2 is connected with the secondary amplification optical path. The utility model is mainly used for spaceflight optical amplifiers.

The invention discloses a carrier rocket hoisting and erecting device and erecting method, belonging to the field of rocket hoisting and launching, comprising a base, a rocket body tail section tray, a hoist and a lifting device; the base comprises a slide steel frame; the slide steel frame has a guide groove arranged vertically and running through it; the upper surface of the slide steel frame is fixedly provided with a guide rail arranged along its length direction; the rocket body tail section tray is detachably fixedly connected with the rocket body tail section; the rocket body tail section tray is provided with a guide wheel group; the guide wheel group comprises guide wheels adapted to the guide rail and arranged in pairs; the hoist comprises a rocket body horizontal hoist and a rocket body flip hoist; the rocket body flip hoist is configured to make the rocket body rotate from a horizontal state to an erected state along the axial direction of the guide wheel group under the action of hoisting; the lifting device is fixedly provided on the rocket body tail section tray. The invention can realize the rapid erection of the rocket by relying on the hoisting device, improve the rocket erection efficiency and shorten the launch time.

A launch system and method for a launch vehicle that can save initial launch fuel includes a support structure, characterized in that: the support structure consists of a horizontal support frame (3) and an ascending section support frame (4), a horizontal track section (5) is installed on the horizontal support frame (3), an inclined track section (6) connected to the horizontal track section (5) is installed on the ascending section support frame (4), a launch acceleration pipe (10) is fixedly installed on the horizontal support frame (3) and the ascending section support frame (4), an electromagnetic catapult (14) is installed on the electromagnetic catapult slide rail (13), a rocket trolley (8) is installed on the horizontal track section (5) and the arc track section (6), and a rocket bracket (9) is provided on the rocket trolley (8). The present invention has a reasonable structure, can obtain sufficient initial velocity before launch ignition, can save launch fuel, can increase the effective payload, and is conducive to saving launch costs.

The invention discloses a high-strength, high-modulus and high-toughness shape memory poly (benzimidazole-benzoxazole) imide material with controllable photodegradability and a preparation method thereof, and belongs to the technical field of multifunctional polymers. The multi-block shape memory poly (benzimidazole-benzoxazole) imide polymer with a photodegradation group is obtained by performing two-step polycondensation and imidization on heteroaromatic diamine containing imidazoles and oxazoles, biphenyl dianhydride and photosensitive group Texamic acid, the tensile strength of the polymer at normal temperature is 262.2 MPa, the Young modulus is 6.0 GPa, the elongation at break is 27.9%, and the polymer can be applied to the field of photodegradation of the polymer. In the shape recovery stage, the energy storage modulus is 100 MPa, the glass transition temperature (Tg) is 380 DEG C, the shape memory performance is achieved, the shape fixing rate is 99%, and the shape recovery rate is 98%. By controlling the proportion of the chemical block photosensitive groups, the polymer has the light-controlled degradation characteristic, and is applied to the fields of light-degradable expandable flexible film space capsules, controllable light-degradable self-adaptive deformation flexible sensors and the like.

The invention provides a movable air source supply device and method for shooting range launching. The movable air source supply device comprises an air supply moving module, a pressurization moving module and an air storage moving module. The gas supply moving module comprises a helium cylinder vehicle and a nitrogen cylinder vehicle; the gas storage moving module comprises a helium gas storage cylinder group and a nitrogen gas storage cylinder group; the pressurization moving module comprises a helium pressurization device and a nitrogen pressurization device, the helium pressurization device is connected with the helium cylinder vehicle and the helium storage cylinder set, and the nitrogen pressurization device is connected with the nitrogen cylinder vehicle and the nitrogen storage cylinder set; the air source system can be integrally transferred and quickly connected; the test flight cost can be reduced, and the preparation period of shooting range launching is shortened.

The application provides a power-oriented electric propulsion gas supply self-learning closed-loop control system and method, which is composed of a storage and supply unit, a power processing unit, a control unit and an electric thruster; the working current of the electric thruster is monitored, if the current is lower than the lower limit, the valve is opened to supplement the gas for a certain time length, the current value of the electric thruster working maximum 3s sliding average within a certain time length is monitored after the gas supplement is completed, and the gas supplement time length is adjusted according to the maximum value with a certain step distance, and the step distance is corrected according to the number of adjustments. The application solves the working condition drift problem of the electric thruster caused by the state change of the spacecraft or the erosion of the thruster itself in the prior art by using the multi-stage self-learning adjustment mode, under the premise of limited on-orbit computing resources, and collects the working parameters for data accumulation and open-loop control of the fault mode.

The invention relates to the field of space remote sensing measurement, in particular to a double-line-scan digital camera optical system based on multiplexing of a primary mirror and a secondary mirror, which comprises the primary mirror, the secondary mirror, a first third mirror, a second third mirror, a rear line-of-sight array focal plane and a front line-of-sight array focal plane, for a rear-view array imaging light path, a light beam is reflected by a first reflection area of a primary mirror, a secondary mirror, a second reflection area of the primary mirror and a first third mirror in sequence and then enters a rear-view array focal plane; for a front-sight array imaging light path, a light beam is reflected by the second reflection area of the primary mirror, the secondary mirror, the first reflection area of the primary mirror and the second third mirror in sequence and then enters a front-sight array focal plane. Through multiplexing of the primary mirror and the secondary mirror, two independent three mirrors are adopted to respectively adjust observation angles corresponding to a double-linear-array imaging light path, the system can flexibly adapt to asymmetric observation angles, and is excellent in image quality, compact in system structure and light in weight under the condition of a large relative aperture. And the method has important engineering significance for promoting the miniaturization and lightweight development of the double-linear-array surveying and mapping camera.

The application relates to a manufacturing method of a head structure of a carbon fiber composite material-based aerodynamic push rod, which comprises the following steps: dividing the head structure into five parts, i.e. a circular truncated cone-shaped upper shell, an inverted circular truncated cone-shaped lower shell, a supporting pipe, a flange seat and a circular truncated cone-shaped protective sleeve; setting a first thickened area and a second thickened area in the upper shell; setting a third thickened area and a fourth thickened area in the lower shell; adopting carbon fiber prepreg to carry out multi-layer pasting on the upper shell, and then inserting multi-layer carbon fiber prepreg into the multi-layer interval of the first thickened area and the second thickened area for pasting; adopting carbon fiber prepreg to carry out multi-layer pasting on the lower shell, and then inserting multi-layer carbon fiber prepreg into the multi-layer interval of the third thickened area and the fourth thickened area for pasting; adopting carbon fiber prepreg to carry out multi-layer pasting on the supporting pipe, and then carrying out curing and forming through a hot pressing tank process after the pasting is completed. According to the head structure requirements, reasonable optimization design of prepreg layer angle, layer sequence and layer proportion is carried out, the head structure bearing requirements are met, and overall weight reduction is realized.

The application discloses an inductance energy storage type pulse power supply with improved energy efficiency and belongs to the technical field of electromagnetic emission. The pulse power supply comprises a primary power supply module, an intermediate energy storage module, a reverse flow loop module, an energy self-recovery module, a load current opening and shutting module and a voltage polarity recovery module; the primary power supply module provides energy for charging the intermediate energy storage module; the reverse flow loop module is used for disconnecting the primary power supply module when the current of the intermediate module reaches a preset value; the energy self-recovery module is used for converting the energy self-recovery of a capacitor; the load current opening and shutting module is used for turning on and shutting off the load current; and the voltage polarity recovery module is used for recovering the voltage polarity of the primary power supply module. The inductance type pulse power supply disclosed by the application has the advantages of high energy efficiency, recycling of inductance and load current tail wave energy, tail wave cutting function, few circuit elements, good circuit expandability, strong repeatability and the like.

The efficient separator comprises a fixed base, a separation connecting piece, a signal collector, a trigger switch and a separation control piece are arranged on the fixed base, the separation connecting piece can be connected with and separated from a satellite fixing platform, and a separation compression spring is connected to the separation connecting piece; when the separation connecting piece is connected with the satellite fixing platform, the satellite fixing platform compresses the separation compression spring, the separation control piece can promote the separation connecting piece to release the connection with the satellite fixing platform, the separation connecting piece can promote the trigger switch to work, and the trigger switch can feed back signals to the signal collector. And the signal collector can feed back information to a ground control system. The device can realize quick response of satellite separation and improve the separation efficiency.

The invention discloses a lightweight design method for a spatial large-aperture reflector based on a Bipod support, and solves the problems that in an existing lightweight design method, lightweight design is independently carried out on a reflector body, and when the Bipod support is installed on the reflector body, the surface shape of the reflector body is affected due to the coupling effect of the Bipod support and the reflector body, and the standard before coupling cannot be met. According to the method, a solid model of the large-aperture reflector is established, then the position of a Bipod supporting boss on the solid model is determined, the lightweight structural form is determined according to the mirror body lightweight rate and the optical surface deformation root-mean-square RMS value of different lightweight structural forms, and then lightweight design is conducted on the solid model of the large-aperture reflector. The influence of the coupling effect between the Bipod support and the large-aperture reflector on the static rigidity and the surface shape of the large-aperture reflector is considered, so that the optical surface deformation index can meet the requirement after the Bipod support is assembled.

The present invention discloses a method for screening the comprehensive radiation shielding performance of materials for deep space exploration, and belongs to the field of radiation protection. This method comprehensively considers the radiation shielding performance of materials in the radiation environment of deep space exploration and the optimal design of shielding materials. The screening process includes determining the radiation environment, constructing a charged particle model that simulates the radiation environment, constructing a geometric structure model of the radiation shielding material, simulating the transport behavior of charged particles in the shielding material, and constructing a database based on the data of multiple radiation shielding material samples to achieve the optimal design of the material. The present invention calculates the influence of the deep space exploration space radiation environment on the radiation shielding performance of the material, obtains the change in the radiation shielding performance of the material system under different structures, and then establishes the theoretical relationship between the radiation environment, the structure of the material, and the performance, so as to provide important theoretical guidance for designing materials with good shielding performance and avoid a large number of trial and error experiments.

The stable rocket launching device comprises a launching base and a launching bin, the launching bin is of a hollow structure, one end of the launching bin is connected with the launching base, a mounting bin is formed in the launching base, the hollow part of the launching bin communicates with the mounting bin, a rocket is arranged in the launching bin, and a power hole is formed in the end, close to the mounting bin, of the rocket; a rotating component is arranged in the power hole, one end of the rotating component is connected with the power hole, the end, away from the power hole, of the rotating component is close to a first primer copper cap in the mounting bin, fuel is arranged in the first primer copper cap, and the end, away from the rotating component, of the first primer copper cap is exposed to the outer side of the launching base. When the stable rocket launching device based on spinning is used, an external firing pin impacts the first ignition copper cap, fuel burns to generate a large amount of high-temperature and high-pressure gas, the rotating component is impacted to rotate, and the rotating component drives a rocket to rotate at a high speed and launch the rocket.

The invention provides a separation spring device on a rocket. The separation spring device comprises an inner sleeve assembly, an outer sleeve assembly, a spring, a first lock hook assembly, a second lock hook assembly and a locking adjusting mechanism. The outer side of the inner sleeve assembly is sleeved with the outer sleeve assembly, and the outer sleeve assembly is movably connected with the inner sleeve assembly; the peripheral sides of the inner sleeve assembly and the outer sleeve assembly are sleeved with the spring, and the first lock hook assembly is installed at the end, connected with the spring, of the inner sleeve assembly and used for connecting one end of the spring with the inner sleeve assembly in a locked mode; the second lock hook assembly is installed at the end, connected with the spring, of the outer sleeve assembly. The locking adjusting mechanism is connected with the inner sleeve assembly and the outer sleeve assembly and used for adjusting relative movement of the inner sleeve assembly and the outer sleeve assembly, adjusting the length of the spring and locking and fixedly connecting the inner sleeve assembly and the outer sleeve assembly. The force value of the spring is adjustable, the spring is detachable, the spring cannot be separated from the first lock hook assembly and the second lock hook assembly, and safe use can be achieved.

The application discloses a multifunctional outer cabin plate dot matrix structure for a satellite, which mainly comprises an inner cabin plate, an outer cabin plate, a side wall plate, a pre-embedded heat pipe, a rod-shaped regular tetrahedron dot matrix structure and a shear-resistant bearing rod, and a satellite single machine is installed on the inner cabin plate, and the outer cabin plate is directly contacted with a vacuum environment and serves as a heat dissipation plate. The pre-embedded heat pipes are all installed on the inner surface of the inner cabin plate, the pipe fin surface of the pre-embedded heat pipe is closely attached to the inner surface of the inner cabin plate, the contact area is increased, and the heat exchange efficiency is improved. The rod-shaped regular tetrahedron dot matrix structure has a whole regular tetrahedron appearance, and the unit cell structure is similar to a diamond single crystal. The rod-shaped regular tetrahedron dot matrix structure is adopted, a plurality of stable triangular structures are formed between the internal supporting arms of the dot matrix unit cell, the mechanical strength and the structural rigidity of the cabin plate are greatly enhanced, and the cabin plate load bearing capacity is improved. Meanwhile, the heat pipes can be pre-embedded in the gap of the dot matrix structure, and low-density and high-thermal-conductivity materials are filled in the blank part, so that the mechanical properties are ensured, and the heat conduction efficiency is improved.

The invention discloses a microwave detonation dual-mode space thruster. The microwave detonation dual-mode space thruster comprises a propellant supply system, a high-thrust propelling module and a micro-thrust propelling module, wherein the high-thrust propelling module and the micro-thrust propelling module are sequentially and coaxially arranged from inside to outside; the high-thrust propelling module comprises an injector, a plasma ignition combustion-supporting device, a detonation chamber, an exhaust nozzle and the like. The micro-thrust propulsion module comprises a microwave supply system, a magnetic pole, a gas collection cavity, a discharge chamber and a grid electrode; the microwave supply system comprises a microwave connector and an antenna; the antenna is a cylindrical antenna, and a convolution cavity is formed between the antenna and the magnetic pole; outer-layer electrons in electrical substance molecules in the convolution cavity do convolution motion in the convolution cavity, and when the convolution angular frequency is equal to the microwave frequency, resonance occurs. The microwave discharge technology, the plasma combustion-supporting technology and the pulse detonation combustion technology are fused, efficient and convenient switching of working modes is achieved through circuit regulation and control, the thruster can meet the strict requirements of large thrust, high specific impulse, long service life and the like, and the thruster can be applied to diversified space tasks.

The embodiment of the present invention discloses a carrier suspension launcher and a carrier suspension launch method, comprising: a rotating base, which is rotatably mounted on a table in a circumferential direction; a pitch adjustment structure, which is arranged on the rotating base and can rotate with an axis parallel to the table as a rotation axis; a track structure, which is detachably connected to the pitch adjustment structure; a positioning support structure, which is detachably positioned and connected to the track structure, and is formed with a mounting portion for placing a carrier to be launched; wherein a lifting assembly is also formed on the pitch adjustment structure, and a connecting assembly that cooperates with the lifting assembly is formed on the track structure, and the lifting assembly and the connecting assembly cooperate with the lifting track structure to move closer to or away from the pitch adjustment structure. The present invention improves the versatility and reliability of the entire launcher by adopting a split arrangement of the track and multiple adjustment structures, and further based on the cooperation of the positioning support structure and the lifting assembly, greatly reduces the preparation time, and reduces the risk of docking operations.

The present invention discloses a space plasma energy collection device and collection method based on the potential difference on the surface of a spacecraft, and belongs to the field of spacecraft energy supply technology. In the present invention, each sensor probe is connected to each of the monitoring points to monitor the potential of the monitoring points; the input end of the processor is connected to the output end of the sensor probe; the output end of the processor is connected to the first input end of the voltage regulator; when the potential of a certain monitoring point reaches the discharge threshold, the monitoring point is turned on to charge the battery using the potential difference between the monitoring point potential and the ground reference point; the monitoring point and the ground reference point are connected to the input end of the inverter; the output end of the inverter is connected to the second input end of the voltage regulator and the battery. The present invention can reduce the accumulation of charge on the surface of the spacecraft, reduce the discharge frequency of the plasma contactor, and has a compact structure.

The embodiment of the present application provides a transfer tool for a launch platform of a carrier rocket, which is used for a launch platform of a liquid carrier rocket. The transfer tool includes: a tool body assembly, which is annular; at least three docking components, which are arranged equidistantly along the circumference of the tool body assembly, are all arranged on one side of the axial direction of the tool body assembly, and are constructed to be detachably connected to the original launch platform; at least three rocket body support seats, plug-in and unplug components, contact feedback components, and charging and discharging connector brackets, are all arranged on the side of the tool body assembly away from the docking components along the axial direction of the tool body assembly; at least three rocket body support seats are arranged equidistantly along the circumference of the tool body assembly. Therefore, when launching different types of carrier rockets, it is only necessary to install the corresponding type of transfer tool on the original carrier rocket launch platform to realize the launch of the carrier rocket, without the need to build multiple carrier rocket launch platforms, thereby shortening the launch cycle in the early stage of carrier rocket development and reducing costs.

The present invention relates to a general platform and test method for fast-response satellite rapid testing based on software definition, which belongs to the technical field of fast-response satellite rapid batch testing, and solves the technical problems of high launch cost, low effective utilization rate, high safety risk, and inability to meet rapid launch requirements of existing fast-response satellites. The solution is: a general platform for fast-response satellite rapid testing based on software definition, including an initial state setting module, a ground equipment comprehensive management module, a first interface adaptation and management module, a second interface adaptation and management module, a bus interaction module, a radio frequency communication interaction module, a platform health check sequence library, a payload health check sequence library, and a result comparison output module; the test method includes the following steps: 1) setting the satellite state; 2) performing satellite-to-ground interface adaptation; 3) channel self-test; 4) platform subsystem health check and payload subsystem health check; 5) result output display. Compared with the prior art, the present invention has the advantages of reduced launch cost of fast-response satellites, high test efficiency, support for efficient and high-intensity continuous testing, and multi-satellite and multi-type parallel rapid testing.