47 results about "Space launch" patented technology

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 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 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.

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 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 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.

The invention belongs to the technical field of spaceflight equipment, and particularly relates to a propellant storage tank and application thereof to a spaceflight vehicle. The invention provides a propellant storage box, the propellant storage box comprises a storage box main body and a bottom-sharing separation layer, the side edge of the bottom-sharing separation layer is connected with the side wall of the storage box main body, and the bottom-sharing separation layer is used for separating the storage box main body in the horizontal direction. The invention further provides an application of the propellant storage tank in a space launch vehicle. In the technical scheme provided by the invention, the common-bottom separation layer replaces an inter-tank section structure, so that the overall weight of the propellant storage tank can be effectively reduced; during preparation, only the common-bottom separation layer needs to be connected with the side wall of the storage tank main body, so that the preparation process is simple and convenient, and the cost is low; the technical defects that in the prior art, a common-bottom propellant storage box is large in weight, high in preparation cost and complex in preparation process are overcome.

The present application relates to a kind of non-throwing aircraft separation launching device, it includes: carrier, the carrier is rotatably mounted with two separation supports, the separation support is opened with gap, two the separation support is locked in closed state by separation bolt, and the cavity is formed between two the separation support, two the gap encloses installation port, the separation support is connected with the carrier, and the elastic member between the carrier tends to open two the separation support;The installation port installs aircraft one, the cavity installs aircraft two.Separation bolt explosion separation, release the limiting fixing of separation support, open separation support under the elastic force of elastic member, release the aircraft one in installation port, also prepare work for subsequent aircraft two separation;Separation support is connected with carrier, retains the rotating function, separation support does not throw after rotating, meet the requirement of not adding space junk to affect orbit safety in space launch activity.

The application provides a method for pre-launch wind correction of a space vehicle, comprising: collecting negative 1-hour measured high-altitude wind data; taking the negative 1-hour high-altitude wind data as an input of trajectory correction, and outputting a guidance parameter; carrying out six-degree-of-freedom Monte Carlo targeting simulation, and calculating flight load; and determining whether to launch the space vehicle according to at least the flight load envelope. The method can improve the launch release probability and reduce the aerodynamic attack angle and flight load.

Implementations of the disclosed subject matter may provide an adaptive thrust structure (400) for a space launch vehicle that may include an interface block (414) that is fluidically coupled to a fuel supply (402), a fuel manifold (406), an oxidizer supply (404), and an oxidizer manifold (408). A first adapter plate (418) may be coupled to a first engine assembly (419). First adapter lines (416) may be fluidically coupled to the first engine assembly (419) and the interface block (414), via the first adapter plate (418). 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 (101), and a reusable and reconfigurable transporter (300) having a plurality of zones that are configurable to transport cargo, one or more passengers, and / or one or more satellites.

The application discloses a global multi-fuel compatible reusable three-stage configuration space-air-ground rapid transport rocket, and belongs to the technical field of space launch vehicles. The rocket adopts a three-stage serial modular layout: the first stage is a multi-fuel adaptive liquid reusable power stage, compatible with multiple liquid propellants, with land and sea dual-base vertical recovery capability, and the reuse frequency is not less than 50 times, and the launch turnaround period is 15-60 minutes; the second stage is a multi-functional on-orbit mission module, suitable for multiple on-orbit functions and large-span loads; and the third stage is a high-energy solid power stage, compatible with traditional and new high-energy solid fuels. The application adopts a quantumization and thermodynamic double correction method to optimize system parameters, and the full-condition calculation error is less than or equal to 1.8%, which can adapt to multiple space missions such as satellite networking, deep space exploration and intercontinental rapid delivery.

The application discloses a spaceflight launching stage wide-frequency random vibration effect simulation method, relates to the technical field of spaceflight engineering and structural dynamics, and comprises the following steps: S001, a multi-physical field coupling model containing thermal load, aerodynamic disturbance and wide-frequency random vibration input is constructed, a variable parameter driving mechanism is adopted to dynamically map material properties and boundary conditions, and a time-varying response curve of structural inherent frequency is acquired; S002, based on the time-varying response curve, modal energy distribution data are loaded, frequency intersection and energy transfer trends are identified, and a response prediction benchmark of a modal overlap region is established. Through thermal-flow-vibration coupling modeling and modal energy analysis, the application realizes early warning and active inhibition of modal jump, forms a closed-loop vibration risk inhibition mechanism through frequency regulation and parameter calibration, significantly improves the response stability and safety of the structure in the launching stage, and has good engineering practicability and innovation.

A system and method are provided for automated pre-launch and post-launch inspection of space launch vehicles and associated launch infrastructure using coordinated swarms of unmanned aerial vehicles (UAVs). In one implementation, multiple autonomous UAVs equipped with multi-modal sensors and protective design features execute coordinated inspection patterns under control of a mission management system that adapts to environmental conditions and enforces safety constraints. Sensor data are processed to detect and classify anomalies, generate inspection reports, and store results for trending analysis and predictive maintenance.

Systems, methods, and devices for dynamically identifying launch opportunities for a space vehicle. A space launch service platform (SLSP) receives time-stamped positional data from an external tracking system monitoring aircraft, maritime vessels, or orbital objects. The SLSP synchronizes received positional data to a unified timeline using a common reference clock. At predetermined intervals, the SLSP calculates predicted future positions of monitored objects by extrapolation. The SLSP calculates predicted proximities between a predefined launch trajectory—defined within an Earth-centered inertial coordinate system—and each predicted future position. The SLSP determines a risk index for each predicted proximity. The SLSP identifies discrete launch windows within a specified launch period where predicted proximities exceed a clearance threshold and associated risk indices remain below a risk threshold. The SLSP aggregates contiguous discrete windows into at least one continuous launch interval and generates a launch opportunity indication associating each interval with a launch readiness signal.

The application provides a leveling force measuring device and system of a space vehicle. The leveling force measuring device comprises: a force bearing module for placing one of the legs of the space vehicle; a force measuring module for measuring the pressure received by one of the legs of the space vehicle; a spherical hinge module arranged between the force bearing module and the force measuring module for transmitting the force in different directions of the space vehicle to the force measuring module; and a lifting module arranged below the force measuring module for participating in the leveling of the space vehicle. The device can meet the test requirements of space vehicles with different diameters by modular combination of the force measuring module, the spherical hinge module and the lifting module, has a compact design structure, and has a few steps of motion transmission between mechanisms and high mechanical efficiency.

The invention discloses a method and a system for predicting simultaneous impact during interval launching of aircrafts, and belongs to the technical field of data prediction processing. The method comprises the following steps: modeling according to an aircraft; changing the initial launching angle and the starting control distance to obtain a launching angle-starting control distance function curve; normalization processing is carried out, and multi-layer image data features are obtained through wavelet decomposition; feature points obtained through wavelet decomposition are used for labeling, and two kinds of wavelets with different parameter coefficients are used for reconstructing an emission angle-starting control distance function; performing difference operation on a function prediction value obtained by prediction and an actual value, and performing multiple rounds of learning to transform parameters of the wavelet itself; reconstructing a prediction function curve by using wavelet transform, and setting miss distance as a trust degree function; and analyzing the obtained wavelet prediction function, setting an initial emission angle and required interval time, and searching a proper interval according to a credibility function. According to the method, the region generated by the minimum value and the maximum value of the overall time can be accurately predicted.

The utility model discloses a rocket erects oil cylinder stop device relates to space launch technology field, and oil cylinder support includes base and is located at the both sides of base's support lug, and support lug is equipped with a plurality of first threaded hole, erects oil cylinder and is equipped with the support of cylinder body and the end of cylinder body, and the support is rotatablely arranged between two support lug through pin shaft and ball hinge bearing, rocket erects oil cylinder stop device includes stop piece and two limit disc. Two limit disc are located at the both sides of support, and limit disc is located between support and support lug, and limit disc is equipped with a plurality of limit hole, and limit hole and first threaded hole one to one correspondence. Stop piece is connected with first threaded hole screw, and stop piece is used for inserting limit hole, makes limit disc and support and holds, to erect oil cylinder and carry out lateral limit, avoid erecting oil cylinder lateral swing. The above-mentioned rocket erects oil cylinder stop device, and erects oil cylinder and can carry out lateral limit through stop piece and limit disc, guarantees erecting frame and erecting oil cylinder can carry out vertical butt joint, and convenient operation.

A multi-launcher space launch assembly includes a movable platform for the assembly, transporting, and vertical positioning of varied launchers. The platform cooperates with a fixed table and a jet deflector, which are situated in a launch area. The infrastructure also includes adaptable fluid refuelling modules compatible with various flow and pressure profiles. The system is adapted to accommodate variations in the height and diameter of launchers by way of modular components.

The invention provides an anti-interference fault-tolerant control method for trajectory tracking of an offshore spaceflight launch and recovery ship, and belongs to the technical field of automatic control of the offshore spaceflight launch and recovery ship. The control method provided by the invention is mainly oriented to research on trajectory tracking control of an offshore spaceflight launch and recovery ship considering unknown external marine environment interference and propeller failure, and applies an interference observer and a fault observer; the problems of online estimation and suppression of unknown external marine environment interference received by trajectory tracking of an offshore spaceflight launch and recovery ship and estimation and compensation of faults of a propeller of the offshore spaceflight launch and recovery ship are solved. The composite anti-interference fault-tolerant controller is used for correcting the tracking feedback control error problem, the anti-interference capability of the marine spaceflight launching and recycling ship is effectively enhanced, the control reliability of the marine spaceflight launching and recycling ship is improved, and the tracking position of the marine spaceflight launching and recycling ship achieves the expected effect.

The utility model discloses a kind of aerospace launching site movable platform electric dual-mode emergency folding device, it is related to movable platform technical field, including fixed frame, the inner wall of fixed frame is fixedly connected with base, the upper surface of base is equipped with the direct-current speed reducer and winch of adaptation, the utility model integrates 220V AC power supply and 48V DC battery dual working mode, when commercial power is stable, rely on 220V AC mode to guarantee efficient operation, when encountering emergency such as sudden power failure, only need manual operation mode change-over switch, system will immediately start 48V DC battery mode, whole process seamlessly links, response is quick, ensure that emergency function continues stable operation, and power coupling of DC motor and AC motor can be realized, accurate control motor speed and torque output, keep stable operating state each time, avoid the phenomenon of shafting caused by uneven stress, prolong the service life of equipment, effectively improve system operation reliability.

A rocket stage for a multistage space launch vehicle, wherein the rocket stage includes a main engine and is configured for stage separation from remaining parts of the space launch vehicle such that the rocket stage returns to the surface of the earth, wherein rocket stage includes an inflatable hull configured to retain lifting gas from a pressure tank, an inflation unit, a propulsion and steering unit configured to provide thrust and attitude control to the rocket stage, and a control unit configured to initiate inflation of the hull by controlling the inflation unit after the stage separation and to control the propulsion and steering unit to maneuver the rocket stage to a specified landing site on the surface of the earth.