27 results about "Space launch" patented technology

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 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 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

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.

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 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 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.

The embodiment of the invention discloses a temperature sensor for seaside spaceflight launching. The temperature sensor comprises a sensor probe and a signal processing circuit. The sensor probe adopts a ceramic matrix with a built-in spring-shaped pure platinum wire, realizes integrated sealing and packaging through a glass glaze layer which is subjected to multiple times of potting and sintering, and has excellent shockproof, waterproof and high and low temperature resistant performances. And the signal processing circuit processes the probe signal based on a pre-stored ultralow-temperature indexing table obtained through liquid helium aging calibration, converts the probe signal into a 4-20mA standard current signal and an RS485 digital signal and outputs the 4-20mA standard current signal and the RS485 digital signal. The core of the invention is to solve the technical problem that the existing sensor cannot realize high-precision and high-reliability measurement and long-distance signal transmission in an ultra-wide temperature range from-258 DEG C to + 100 DEG C, and is especially suitable for ultralow temperature monitoring in severe environments such as a seaside spaceflight launching site.

The invention relates to the technical field of fuel waste gas, in particular to an aerospace launch hydrazine fuel waste gas purification treatment device and method.The aerospace launch hydrazine fuel waste gas purification treatment device comprises a plurality of tubular reactors connected in series, the inner walls of the tubular reactors are coated with titanium-based catalyst coatings, the upper portions in the tubular reactors are oxidation reaction chambers, and the lower portions in the tubular reactors are ozone treatment chambers; an ozone aeration disc is mounted at the bottom in the ozone treatment chamber, an ozone inlet is formed in one side of the ozone aeration disc, the ozone inlet is formed in the side wall of the ozone treatment chamber, and a granular porous adsorption material is laid on the top surface of the ozone aeration disc; an ultraviolet lamp tube is arranged between the top in the tubular reactor and the top surface of the granular porous adsorption material, and the ultraviolet lamp tube is mounted at the top in the oxidation reaction chamber; the bottom of the tubular reactor is communicated with a drainage pipe. The device is not only suitable for treating waste gas generated in the hydrazine fuel transfer and filling process, but also suitable for hydrazine fuel production and transportation and purification treatment of hydrazine fuel waste gas in other places.

This invention proposes a reusable waterborne takeoff and landing space launch vehicle, belonging to the field of aerospace technology. It includes a rocket body, wing surfaces, tail fins, propulsion engines, attitude control engines, floats, mounting brackets, and a deceleration parachute release device. The rocket body has a pointed cylindrical shape and houses propellant, oxidizer, navigation, control, and energy supply equipment. Wing surfaces are mounted in pairs on both sides of the rocket body. Floats are symmetrically mounted on the wing surfaces on both sides. Tail fins are mounted in pairs on both sides of the tail section of the rocket body, and control surfaces rotating around the wing surfaces are mounted on the tail fins. A propulsion engine, consisting of a set of rocket engines and / or aero engines, is mounted at the end of the rocket body. A set of attitude control engines is symmetrically arranged at one end of the rocket body. The mounting bracket is axially positioned on the upper side of the rocket body. The deceleration parachute release device is mounted on the upper side of the rocket body and houses the deceleration parachute. This invention also proposes a method of use, solving the problems of high space launch costs and difficulty in increasing the maximum payload capacity.

The utility model 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 utility model, 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.

The application is suitable for the field of space launch technology, and provides a ground gas supply and distribution system capable of realizing automatic self-checking, comprising a control system and valve action test gas path modules, pressure reducing valve steady pressure performance test gas path modules, air tightness test gas path modules and pressure transmitter calibration gas path modules electrically connected with the control system; the valve action test gas path modules and the air tightness test gas path modules have the same gas path structure, and share the same set of hardware gas paths; the control system realizes valve action test and air tightness test functions on the hardware gas paths by executing different automatic test processes; the application simplifies the system gas path structure, realizes automatic self-checking of the whole process, greatly reduces the operation intensity of personnel, shortens the test process, reduces human error operation, and improves the system test efficiency and reliability.

The invention provides a thunder and lightning protection system for a spaceflight launching site, and the system comprises four lightning protection towers which are equal in height and are used for installing lightning rods and lightning arresting lines, and transmitting the thunder and lightning current striking the lightning rods and the lightning arresting lines to a grounding grid of the launching site; the four lightning rods are the same in specification, are mounted at the top ends of the four lightning protection towers respectively, and are used as lightning arresting devices to conduct attracted lightning current to the lightning protection towers; the lightning arresting lines are arranged among the four lightning arresting towers in three layers, and the lightning arresting lines on each layer are connected in a rectangular manner by taking the four lightning arresting towers as vertexes; the four lightning protection towers, the four lightning rods and the lightning arresting lines form a cage-shaped three-dimensional rectangular lightning protection area. In the technical scheme provided by the invention, the cage-shaped three-dimensional thunder and lightning protection area can play a good thunder and lightning protection effect on the top and the side surfaces of the protected object. Moreover, the lightning rod and the lightning arresting line are used as lightning arresting devices, so that double or multiple protection is provided for the protected object, and the reliability of lightning protection is improved.

This invention provides a performance verification method for a test block of a material for repairing aerospace launch site guideways, belonging to the field of material performance verification. The method includes: constructing a verification database based on performance verification-related data; constructing and training a performance prediction model based on the verification database; inputting the attribute data and target operating condition data of the test block to be verified into the performance prediction model to obtain initial performance prediction results; designing and executing physical calibration tests based on the initial performance prediction results to obtain test results; comparing the initial performance prediction results and the test results to determine whether the performance prediction model has passed calibration verification; if the performance prediction model has passed calibration verification, the initial performance prediction results are used as the performance verification results of the test block to be verified. This invention, by virtually predicting material properties and supplementing it with a minimal amount of targeted physical calibration, can reduce reliance on high-cost destructive testing and significantly shorten the verification cycle.

The invention relates to an interface (4) for securing a load to be placed in orbit to a launch platform, the interface comprising two plates (10, 12) borne by the platform and by the load, respectively. The plates can be secured to one another. A first plate bears at least two locking elements (14) which are movable between two states. In the approach state, the plates can be brought together without being locked by the locking elements. In the locked state, the plates are secured together, a bolt of each locking element engaging with a keep on the second plate. The locking elements can adopt an intermediate, captive state, in which they engage with the second plate while allowing translational and rotational movements, within predetermined ranges, thereof. The transition from the captive state to the locked state gradually causes the plates to rotate and translate into the secured position.

The invention relates to an exhaust nozzle (1) of a space launcher engine equipped with a cooling system, extending along a longitudinal direction (D) and having a proximal end intended to be connected to a combustion chamber and a distal end (10), the cooling system comprising a lower manifold (20) disposed at the distal end and along the periphery of the nozzle and channels (30) extending along the longitudinal direction communicating with the lower manifold, characterized in that the cross-section of the lower manifold decreases in at least one dimension along the length of the manifold. Figure to be published with the abstract: Fig. 1