85 results about "Space launch" patented technology

Systems and methods for providing a space launch service platform (SLSP) that integrates artificial intelligence and data analytics to support launch operations. The SLSP may connect to multiple data sources, collect data, and standardize the collected data according to regulatory and operational standards. The SLSP may evaluate launch safety and risks using standardized data and generate a situational analysis for decision-making. The SLSP may provide graphical overlays and decision-support tools to highlight optimal launch windows and potential risks.

The invention belongs to the field of complex equipment system design and verification, and discloses a model-based complex equipment system design and verification integrated method, which takes four layers of'demand-behavior-structure-parameter 'as a framework, combines task scene demands, and constructs a full-process design and analysis method of a reusable space launch vehicle system. The method comprises the following steps: decomposing a complex system from demand analysis, performing parameter calibration and performance verification on a model by combining theoretical and experimental data, constructing a reusable space launch vehicle system model, and performing comprehensive analysis and verification of the system; in combination with a three-dimensional virtual environment and a configuration optimization algorithm of the model, performing rapid optimization on various configurations of the aerospace vehicle, and evaluating feasibility and performance of a design scheme; through bidirectional interaction of design and verification models and communication of system-level design and professional-level design, the problems of design and verification separation, insufficient model collaboration and the like in the prior art are solved, and the efficiency and precision of design and verification of a complex equipment system are improved.

Implementations of the disclosed subject matter may provide an adaptive thrust structure for a space launch vehicle that may include an interface block that is fluidically coupled to a fuel supply, a fuel manifold, an oxidizer supply, and an oxidizer manifold. A first adapter plate may be coupled to a first engine assembly. First adapter lines may be fluidically coupled to the first engine assembly and the interface block, via the first adapter plate. A second engine assembly, second adapter lines, and a second adapter plate are capable of replacing the first engine assembly, first adapter lines, and first adapter plate. Implementations of the disclosed subject matter may provide a reusable and reconfigurable launch vehicle that may include a reusable second stage, and a reusable and reconfigurable transporter having a plurality of zones that are configurable to transport cargo, one or more passengers, and / or one or more satellites.

The invention discloses a fixing device for a rocket, and relates to the technical field of spaceflight launching equipment, and the device comprises a bearing mechanism which is provided with an embracing clamp assembly and an embracing clamp adjusting assembly for driving the embracing clamp assembly to act; the holding clamp assembly comprises two holding clamp main bodies, each holding clamp main body is provided with a connecting section and a clamping section, the holding clamp adjusting assembly comprises a first mounting seat, a second mounting seat and a limiting seat which are mounted on the bearing mechanism, a driving part is arranged on the first mounting seat, a driving screw rod is rotationally connected between the second mounting seat and the limiting seat, and a driving shaft of the driving part is connected with the driving screw rod; the driving lead screw is provided with a first lead screw section and a second lead screw section which are oppositely arranged in a threaded mode, the driving lead screw is driven by the driving piece to rotate, so that the first driving nut and the second driving nut move along the driving lead screw, then the two holding clamp bodies are driven to rotate relative to the bearing mechanism, and the two clamping sections are matched to clamp or loosen a rocket to be fixed. And the operation process is simple and efficient.

The invention provides a spaceflight launch mission evaluation method and system, and the method comprises the steps: dividing a spaceflight launch mission into a plurality of rocket flight segments, and carrying out the evaluation of a subsystem level, a rocket level and a key action in each rocket flight segment. Performing task level evaluation by combining the previous evaluation result, the parameter value of the injection parameter and the like after satellite-rocket separation to obtain a final task evaluation result of the spaceflight launch task; wherein the evaluation of the evaluation elements in the subsystems and the evaluation of the subsystems are realized by using a process evaluation criterion, and the evaluation of the subsystem level, the rocket level and the key action is realized by using a result evaluation criterion. According to the scheme disclosed by the invention, each level can take the most direct and most critical result index as an evaluation basis from the whole to the part, so that subsequent improvement and optimization are effectively guided.

Flexible hole forming equipment suitable for a rocket tube section end frame comprises an axial hole forming device, a radial hole forming device, a rotary table, a flexible tool and a control system. Wherein the rotary table is fixed on a foundation through an adjusting sizing block and a foundation bolt; the flexible tool is mounted on the rotary table through a positioning pin and a screw; the axial hole forming device and the radial hole forming device are installed on the two sides of the rotary table at the interval of 90-degree azimuth angles. The device is mainly used for machining axial butt joint holes and radial riveting holes of multi-model cylinder section end frames of spaceflight carrier rockets and axial milling, flexible hole forming of the rocket cylinder section end frames with the diameter of phi 3350 mm and phi 3800 mm can be met, and machining of all inner L-shaped and outer L-shaped end frames with the diameter of phi 3350 mm and phi 3800 mm is considered; the problem that a multi-drill jig tool corresponding to multi-type end frames is tedious to replace is solved, the size precision and the position precision of end frame drilling are greatly improved, and the machining period is shortened.

The present invention relates to the technical field of electronic digital data processing, and in particular to a training and evaluation system for a space launch site. The system comprises an instructor platform, an integrated platform for simulation training devices, and a video monitoring platform. The instructor platform includes an integrated training scenario setting module, an instruction module, and an evaluation module. In this solution, task parameters and process logic are customized through the instructor platform, the simulation training device provides highly simulated equipment operation feedback, and the video monitoring platform collects behavioral data without blind spots, forming an integrated "training-execution-evaluation" management and control system. Based on a time node partitioning algorithm, an equipment state vector model, and a weighted average evaluation algorithm, the system quantitatively analyzes trainees' operational accuracy (such as mean square error compared to equipment status), timeliness (task node completion time deviation), and collaboration ability (task allocation matrix tracking role performance), thereby addressing the subjectivity and lag issues of traditional evaluation.

The satellite supporting cabin comprises an upper end frame, a truncated cone shell and a lower end annular flange which are integrally formed by carbon fiber composite materials, the upper end frame is formed by wrapping a first foam sandwich with the carbon fiber composite materials and solidifying the first foam sandwich, and the upper end frame is used for being connected with a satellite support. The lower end annular flange is used for being connected with a substage of an aerospace vehicle. The satellite supporting cabin can achieve the purpose of reducing the weight and has high rigidity.

This invention belongs to the field of space laser communication and deep space exploration technology, and discloses an integrated detection device for space laser communication monitoring and interaction. The device includes a main body, on which are integrated a solar panel, a navigation imaging device, a communication antenna, a communication receiver, a radar, a landing support, and functional components for communication monitoring, environmental perception, and interactive operation. This invention forms a closed-loop system through electrical, signal, and mechanical connections, enabling simultaneous execution of multiple tasks, significantly improving the device's space operation efficiency, reducing the load pressure on space launch vehicles, and adapting to the space constraints of on-orbit operations. The communication antenna and receiver achieve bidirectional signal interaction through a high-speed laser communication module and are linked with the navigation imaging device and environmental sensors to ensure stable transmission and status monitoring of laser communication signals, which is beneficial for practical applications and operations.

The invention relates to the technical field of spaceflight launching equipment testing, in particular to a load test device and method for a fixed launching pad of a medium-large carrier rocket. The technical problems that in the prior art, a fixed launching pad load test device is insufficient in bearing capacity and single in loading mode, and multi-point loading and ballasting working condition tests cannot be met at the same time are solved. Comprising a test platform, a test tool, loading equipment and test equipment. The test tool comprises a base, a portal, a loading beam, a support and a lug, and is connected with a foundation embedded steel plate through bolts; the loading equipment consists of a plurality of hydraulic jacks and a hydraulic pump station, and is uniformly controlled by a PLC (Programmable Logic Controller) multi-point synchronous jacking control system; and the test equipment comprises a micrometer and a laser tracker and is used for accurately measuring the deformation of the fixed table body. The device is simple in structure, safe and reliable in structure, flexible in loading mode, high in bearing capacity and capable of achieving multi-point-position and multi-working-condition load tests with the maximum 2025 t ballast and 625 t pull load.

The utility model discloses a spaceflight launching and recycling situation visualization system, which belongs to the technical field of visualization systems and comprises a data acquisition module, a data server cluster, an animation server, a large-screen display device and a switch. The data server cluster is in communication connection with the switch, the switch is in communication connection with the animation server, and the data server cluster is in communication connection with the large-screen display device. According to the method, a data server cluster is established, a topographic map of a recovery position is made, and real-time three-dimensional animation production is performed through data obtained from the data server cluster when a spacecraft is launched and recovered, so that visualization when the spacecraft is launched and recovered can be realized, an intuitive space display mode can be realized, and understanding is facilitated.

The invention relates to a multi-source electromagnetic matrix aircraft take-off and landing system and a waste heat utilization and building integration method thereof, and belongs to the technical field of aircraft take-off and landing. The system comprises at least three electromagnetic emission units which are arranged in a space array, and each electromagnetic emission unit comprises an independent controllable excitation coil; the central controller is used for independently adjusting current parameters of each electromagnetic emission unit; the at least three electromagnetic emission units act together to form at least four independent and controllable electromagnetic stress action points in a predetermined space area. The system comprises a building facade integration system, a waste heat recovery system and a military application system (space debris cleaning and aerospace vehicle launching). The system has the multiple advantages of high control dimension, high magnetic field intensity, low cost, mobility, building embeddability, comprehensive energy utilization, military and civilian dual-purpose, military defense value and the like, and can be used as a general take-off and landing infrastructure in the fields of urban air traffic, military aviation, spaceflight launching, military defense and the like.

The invention belongs to the technical field of underwater robots, and particularly relates to an underwater robot preset performance flexible interference compensation control method. The control method mainly aims at the research of underwater robot trajectory tracking control considering unknown external marine environment interference, and uses an interference observer, a preset performance control technology and a flexible interference compensation technology. The problems of online estimation, suppression and compensation of unknown external marine environment interference on trajectory tracking of the underwater robot and tracking control under preset performance constraints are solved. A preset performance flexible interference compensation controller of the underwater robot is used for correcting a tracking feedback control error problem, so that the anti-interference capability of the underwater robot is effectively enhanced, the reliability of marine spaceflight launching and recovery ship control is improved, and the tracking position of the underwater robot achieves an expected effect under the constraint of user-defined performance requirements.

The present invention belongs to the technical field of refractories, and particularly relates to a material resistant to ultra-high temperature and high-speed jets for vehicle-mounted deflectors at space launch sites and a construction process thereof. In the present invention, components such as mullite and corundum are selected as matrix materials, while nano-scale superfine powders such as nano-silicon carbide and nano-aluminum oxide are added. In addition, a small amount of graphite is added for adjustment to obtain a material resistant to ultra-high temperature and high-speed jets, effectively improving the normal temperature compressive strength of the material, reducing the ablation rate, and simultaneously obtaining good refractoriness.

The invention relates to the technical field of spaceflight launching, and provides a data processing method and system for manned spaceflight launching, electronic equipment and a medium, and the method comprises the steps: S1, obtaining a process item set of a manned spaceflight on-duty launching technology; s2, acquiring four modes of manned spaceflight on-duty launching; s3, setting a security recovery point; s4, setting emission technology process item sets corresponding to the four modes; s5, acquiring a current transmitting condition; if the emergency rescue launching condition is met, launching is carried out according to an emergency rescue technical process item set; s6, if the conventional rescue launching conditions are met, launching is carried out according to a conventional rescue technology project set; s7, in the implementation process, if emergency rescue launching conditions are met, launching is carried out according to an emergency rescue technology project set; if conventional rescue launching conditions are met, launching is implemented according to conventional rescue technology and process projects; and S8, repeating the steps S5 to S7 until the emission is completed. According to the scheme, on the premise that reliability and safety are guaranteed, the task process cycle is shortest.

The invention provides a space docking interface based on electromagnetic driving, and relates to the technical field of space launching technologies. Comprising a child star interface and a mother star interface, the child star interface comprises a child star interface body, a child star coil and a coil iron core are installed on the child star interface body in an embedded mode, and a locking rod extending outwards is installed on the lower portion of the child star interface body; the mother star interface comprises a mother star interface main body, a mother star coil is embedded in the mother star interface main body, and a locking assembly is arranged at the upper part of the mother star interface main body; the child star coil and the mother star coil are electrified, the lower end face of the child star interface makes contact with the upper end face of the mother star interface, and the locking rod is matched with the locking assembly. A locking scheme of an electromagnetic-locking assembly is adopted, so that fuel consumption and plume pollution are avoided; the current in the primary star coil is in a dynamic adjustment stage, the current of the secondary star coil is gradually reduced along with the reduction of the distance between the primary star and the secondary star, and the flexible butt joint is realized by good buffer performance; the pre-tightening rod is pressed downwards to drive a contact of the microswitch to act, and a locking signal is sent out while pre-tightening is achieved.

The present application provides a space launch mission path planning method, device, electronic device and medium, the method comprising: extracting information from geographic information corresponding to the space launch mission and constructing a geographic topology relationship graph; splitting the geographic topology relationship graph according to the hierarchical relationship between different nodes to obtain multiple geographic topology relationship subgraphs; determining multiple candidate planning paths in each geographic topology relationship subgraph, and screening the multiple candidate planning paths based on the space launch mission performance parameters to determine the local planning path corresponding to the geographic topology relationship subgraph, connecting the first local planning path with the second local planning path to generate the best global planning path. By adopting the above-mentioned space launch mission path planning method, device, electronic device and medium, the problem of long search time of the path planning algorithm when the road network scale is large, the search space and the complexity of the space launch mission are high is solved.

The invention relates to the technical field of spaceflight launching engineering, in particular to a whole solid carrier rocket transshipment method in a specific limited area, which comprises the following steps: acquiring the mass and hoisting attribute of a to-be-transshipped solid carrier rocket and the scale condition data of an existing launching site; designing a transshipment method thought according to the obtained to-be-transshipped solid carrier rocket and the launching information of the existing launching site, and determining a transformation technical scheme of the existing launching site; designing a carrier rocket transshipment test launch process, and implementing integral hoisting of the carrier rocket. By designing a transshipment scheme of module flat car rail extension and gantry crane and truck crane cooperative hoisting, safe and efficient transshipment of the whole solid carrier rocket under the condition of a limited launching site is achieved, and the stability of transshipment operation and the site space utilization rate are remarkably improved; and an economical and feasible technical path is provided for an old launching site to adapt to novel rocket guarantee requirements.

A thermal environment deflection test method for electronic products is used in deflection testing of single units supporting space launch vehicles and missile weapon systems. This method is used to determine product design margins, identify weaknesses, and conduct quantitative deflection testing to improve product reliability. This invention is a reliability enhancement method based on quantitative deflection testing.

The invention discloses a carrier rocket trajectory simulation method, system and device and a medium, and relates to the technical field of aerospace vehicle trajectory design, and the method comprises the steps: carrying out the trajectory parameter initialization according to the overall parameters of a carrier rocket and a flight program instruction, and obtaining the initial simulation time, the initial simulation step size and the initial rocket state parameters; determining a rocket state parameter corresponding to the next simulation moment according to the initial simulation time, the initial simulation step length and the initial rocket state parameter; determining a target state parameter based on the rocket state parameter corresponding to the next simulation moment and a judgment result of a preset condition, and outputting feature trajectory data based on the target state parameter; and obtaining compressed carrier rocket trajectory simulation result data based on the characteristic trajectory data. According to the method, carrier rocket trajectory feature point parameters can be accurately captured, output data is compressed, and the simulation calculation efficiency is improved.

The invention is suitable for the field of spaceflight carrier rockets, and provides a vertical recovery carrier rocket propellant storage box which comprises an upper cabin, a lower cabin, a partition plate assembly, a flow accumulator, an anti-shaking plate and an anti-vortex and anti-collapse device. The storage box is divided into an upper cabin and a lower cabin through the partition plate assembly. The lower cabin is provided with a flow accumulator, and an outlet of the flow accumulator is provided with an anti-vortex and anti-collapse device; and the anti-shaking plates are unevenly distributed on the upper cabin and the lower cabin along the axis of the storage box. According to the invention, the management problem of vertical recovery carrier rocket propellants is solved, and the ignition condition of multiple starting of an engine can be met; a special bottom sinking system is omitted, and compared with an existing scheme adopting a bottom sinking engine, the system complexity is reduced, the mass of a rocket body is reduced, and the reliability of first-sublevel recycling work is improved.

The invention relates to the technical field of spaceflight carrier rockets, and provides a large engine high modulus test vacuum chamber gate which at least comprises a gate body. The hanging foundation comprises a stand column and a cross beam; a connecting part of the lifting mechanism is connected with the stand column, and an executing part of the lifting mechanism is connected with the cross beam; the moving mechanism is arranged on the cross beam; one end of the hanging mechanism is connected with the moving mechanism, the other end of the hanging mechanism is connected with the door body, and the door body is driven by the moving mechanism to move in the axial direction of the cross beam; the door groove is formed below the door body; the groove filling mechanism is arranged on the side, away from the vacuum chamber, of the door groove. According to the large engine high modulus test vacuum chamber gate, the supporting and moving problems of the large-size and large-mass vacuum chamber gate are solved, and the problem of clamping stagnation of opening and closing of the gate due to excessive deformation of the rotating shaft is solved. And moreover, a track does not need to be arranged on the ground, the groove filling mechanism can flatten the door groove after the door body is opened, and entering and exiting of the engine tool car are prevented from being affected.

The application provides a kind of offshore aerospace launch and recovery ship power positioning fault-tolerant anti-interference method, belongs to offshore aerospace launch and recovery ship automatic control technical field;The method of the application is mainly oriented to the research of offshore aerospace launch and recovery ship power positioning control considering unknown external marine environment interference and mixed faults of thrusters, uses interference observer and bias fault observer and multiplicative fault adaptive observer, solves the online estimation and inhibition of unknown external marine environment interference on offshore aerospace launch and recovery ship power positioning and the estimation and compensation problem of mixed faults of offshore aerospace launch and recovery ship thrusters;By using a composite fault-tolerant anti-interference controller to correct the tracking feedback control error problem, effectively enhance the anti-interference ability of offshore aerospace launch and recovery ship, improve the reliability of offshore aerospace launch and recovery ship control, so that the expected position of offshore aerospace launch and recovery ship reaches the expected effect.

The utility model discloses a spaceflight launching equipment structure crack detection device which comprises a hydraulic lifting system, a control unit, an ultrasonic flaw detection probe, a conductive slip ring, a cable, a data acquisition assembly and a digital twin system. The data acquisition assembly is used for acquiring images of the surface of the structure, the control unit is used for transmitting information detected by the ultrasonic flaw detection probe and acquired by the data acquisition assembly to the digital twin system, and when crack loss exists in the structure, the position and degree can be visually displayed, positioning is facilitated, timely processing is facilitated, and the working efficiency is improved. And during detection, automatic control is facilitated, and use is convenient.

The present invention relates to a space launch vehicle for transporting a load from the earth to a space orbit or to space and, more specifically, to a carbon composite ultralight space launch vehicle fuselage manufactured using a carbon fiber reinforced composite for the fuselage of the launch vehicle and a propellant tank accommodated in the fuselage, and a manufacturing method thereof.

The present application relates to the technical field of electric data processing, and especially relates to a multi-launch site space launch task planning and scheduling method based on resource residual quantity, which comprises the following steps: S1, constructing a launch task set based on launch tasks and constructing an available resource set based on existing resources; S2, constructing a resource requirement set for each launch task; S3, according to the launch task set, the available resource set and the resource requirement set, determining a global optimal solution of a planning and scheduling model based on an improved ant colony algorithm, and performing launch task planning and available resource scheduling according to the global optimal solution. The present application fully considers the relationship among launch tasks, task requirements and available resources, and can improve the reliability of space launch task planning and scheduling.

The invention relates to a method for initiating or igniting a mixture of an oxidizer and a fuel in a combustion chamber of a space launcher, characterized in that it comprises implementing a step of injecting decomposed hydrogen peroxide (HTP) via a catalyst so as to obtain a predetermined ignition temperature. Figure to be published with the abstract: Fig. 1

The invention provides a falling object prevention and control device used in test operation of a spaceflight launch site, belongs to the field of falling object prevention and control matched with test use, and solves the problem of unmanned falling object prevention and control in the test operation of the spaceflight launch site. Comprising a locking piece, a stretching piece, a carrying bag and a rotating hinge. The outer side of the bottom of the locking piece is connected with the rotating hinge; the rotating hinge is connected with the opening part of the carrying bag through the hinges at the two ends of the stretching piece; the locking piece is clamped and fixed to a fixed part of a spaceflight launching site, the angle and position of the stretching piece are adjusted after the rotating hinge connected with the outer side of the bottom of the locking piece rotates, the stretching piece adjusts the position of the object carrying bag through the hinges at the two ends, and the object carrying bag is made to be arranged below a target falling object. According to the invention, unmanned anti-falling measures in the test operation of a spaceflight launching site are realized, the quality safety requirements of falling object prevention and control in the test operation are met, and the safety and reliability in the test operation process are improved.

This invention provides an electromagnetically driven space docking interface, relating to the field of space launch technology. It includes a satellite interface and a mother satellite interface. The satellite interface includes a main body with a satellite coil and coil core embedded in it, and an extended locking rod at the lower part of the main body. The mother satellite interface includes a main body with a mother satellite coil embedded in it, and a locking assembly at the upper part of the main body. Both the satellite and mother satellite coils are energized, with the lower end face of the satellite interface contacting the upper end face of the mother satellite interface, and the locking rod engaging with the locking assembly. This electromagnetic-locking assembly locking scheme results in no fuel consumption and no plume pollution. The current in the mother satellite coil is dynamically adjusted; as the distance between the satellite and mother satellite decreases, the current gradually decreases, providing excellent buffering performance and enabling flexible docking. The pre-tightening rod presses down, actuating the contacts of a microswitch, simultaneously achieving pre-tightening and sending a locking signal.

1.The name of the design product: space launch control hall launch cabin. 2.The use of the design product: simulate the astronauts in the process of rocket launch weightlessness. 3.The design points of the design product: the combination of shape, pattern and color. 4.The picture or photo that best shows the design points: front view. 5.The design claimed contains color.