15 results about "Launch window" patented technology

This invention discloses a method, apparatus, device, and storage medium for satellite-rocket coordinated mission planning. The method determines constellation configuration parameters based on the performance requirements of the current satellite-rocket coordinated mission and satellite resources. The constellation is then divided into several subsets based on these parameters, with each subset corresponding to a certain number of launch missions. Based on the number of launch vehicles and support positions at each launch site, each subset is allocated to each launch site according to a preset ratio. A preliminary deployment plan is generated based on the launch vehicle selection, launch site planning, and launch window planning at each launch site. The preliminary deployment plan is then used to verify whether the current launch resources meet the constraints. If they do, the current satellite-rocket coordinated mission is executed; otherwise, the constellation configuration parameters are redefined. This method achieves efficient coordination in satellite-rocket coordinated mission planning through a dynamic closed-loop optimization mechanism, improving resource utilization and ensuring mission feasibility under timeliness, launch window, and support position constraints.

The invention discloses a two-way rocket launching window searching method, system and device and a medium, and the method comprises the steps: rapidly calculating two nearest launching windows of an ascending orbit and a descending orbit of a rocket according to the attributes of a central body, target orbit plane parameters, the geographic coordinates of launching points and the flight duration of the rocket; according to the method, the problem is converted into numerical calculation based on space geometry, the crossing moment is directly solved by using the geometrical relationship between the normal vector of the orbital plane and the inertial position vector of the launching point location and adopting efficient numerical methods such as dichotomy and the like, and complex dynamic iteration is avoided. According to the method, the calculated amount is greatly reduced, rapid searching of the launching window is achieved, the method is suitable for an interstellar detection task scene needing rapid response or multiple times of calculation, and the real-time performance of task planning is remarkably improved.

The application discloses a multi-shot window breaking device, and relates to the technical field of window breaking devices.The device comprises a window breaking main shell, a window breaking assembly and a control assembly.The top of the window breaking main shell is used for being carried by a drone device.The window breaking assembly is used for launching window breaking bullets to a target window, and comprises a plurality of arranged launching barrels.Each launching barrel is connected with the window breaking main shell, and the inside of each launching barrel is provided with a window breaking bullet.The launching barrel is used for limiting the launching path of the window breaking bullet.The control assembly is in signal communication with the window breaking assembly and the drone device, and is used for cooperating with the drone device to control the window breaking assembly to launch window breaking bullets to the target window.The device can quickly assemble the window breaking assembly, thereby effectively improving the efficiency of the multi-shot window breaking device when performing a window breaking task.

The present invention discloses a method and system for optimizing the design of Earth-Moon orbits under a four-pulse ephemeris model based on manifold splicing, and relates to the technical field of designing probe transfers between special Earth-Moon orbits. The technical key points of the present invention include: processing the initial values ​​of the Halo orbits corresponding to the Sun-Earth system and the Earth-Moon system based on invariant manifolds, and obtaining the optimal manifold splicing points of the Halo orbits corresponding to the Sun-Earth system and the Earth-Moon system respectively; constructing a four-pulse dual-circle restrictive four-body model; using the optimal manifold splicing points of the Halo orbits corresponding to the Sun-Earth system and the Earth-Moon system as initial values, and using a particle swarm optimization algorithm and a sequential quadratic programming algorithm to solve the four-pulse dual-circle restrictive four-body model to obtain the optimal spliced ​​orbit. The present invention uses two optimization models to design low-energy transfer orbits, which can accelerate iterations, improve computational efficiency and algorithm convergence, and realize rapid search for low-energy transfer orbits in specific launch windows, meeting the needs of space stations for Earth-Moon low-energy orbit transfers.

The invention discloses a vertical launch vehicle launch window prediction control method, system and device and a medium, and relates to the technical field of unmanned vehicles. The method comprises the steps that the six-degree-of-freedom motion state of the vehicle is sensed in real time through a multi-source sensor; the six-degree-of-freedom motion state comprises a three-dimensional position, a three-dimensional posture and corresponding linear velocity and angular velocity; forecasting the future short-time movement track of the vehicle based on the wave dynamics model; judging whether a dynamic launching window meeting the safe launching condition of the unmanned aerial vehicle exists or not according to the forecast result; based on the six-degree-of-freedom motion state, the attitude of the vehicle is controlled through cooperation of multiple execution mechanisms, so that the vehicle enters and is stabilized in the launching window; and triggering an unmanned aerial vehicle launching instruction after confirming that the attitude of the vehicle is stable. In a semi-floating state, the attitude of the unmanned aerial vehicle can be sensed, predicted and actively controlled, and the optimal opportunity is created and captured for vertical launching of the unmanned aerial vehicle in the cabin.

The application provides a spacecraft orbit transfer section attitude conversion time sequence generation method and system, including the following steps: step S1: obtaining a satellite multiple orbit transfer time sequence; step S2: determining a satellite attitude conversion time during an orbit transfer section according to the satellite multiple orbit transfer time sequence; step S3: compiling an attitude conversion semantic matrix according to the attitude conversion time; step S4: converting the attitude conversion semantic matrix into a real attitude time sequence matrix; and step S5: checking the real attitude time sequence matrix. The application uses a simple matrix to represent an attitude time section, and further converts the attitude time section into a real attitude time section through logical judgment. The generation of the attitude time sequence can be quickly realized by computer programming, and the application has strong practicability in launch window calculation and the like.

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 invention discloses a spacecraft rendezvous and docking method and device based on step-by-step orbit lifting phase modulation control, and relates to the technical field of space.The spacecraft rendezvous and docking method comprises the steps that on the basis of a preset launching window range, the coplanar phase difference between a target spacecraft corresponding to each launching window and a visiting spacecraft is determined, and based on the determined date, the determined phase error reference, the coplanar phase difference, the phase error reference and the target phase difference, determining a target launching window in a preset launching window range, and based on a step-by-step orbit lifting phase modulation control strategy, launching the visiting spacecraft in the target launching window. According to the invention, the technical problem of low phase modulation control efficiency caused by incapability of accurately determining the launching window in a spacecraft rendezvous and docking task in the related technology is solved.

The invention relates to a carrier rocket flight path online planning method considering collision constraints. The method comprises the steps that a flight path online planning dynamic model of a carrier rocket is established; setting a collision constraint condition based on the flight path online planning dynamic model, and solving an orbit injection state of the carrier rocket in avoiding a collision area based on a sequential quadratic programming algorithm; based on the current speed and position information of the carrier rocket, whether orbit degradation is needed or not is judged; and if yes, designing a degraded track. According to the carrier rocket flight path online planning method considering the collision constraint provided by the invention, flight path online planning is carried out based on the real-time speed and position information of the carrier rocket, the collision constraint is considered, and the collision risk with a space object is calculated online, so that a collision area automatic avoidance function is realized, and available launching windows are effectively expanded; and meanwhile, whether orbit degradation is carried out or not is decided, so that on-line re-planning and high-reliability orbit injection of the flight path of the carrier rocket are realized, and hysteresis of traditional off-line planning is avoided.

This invention provides a method, apparatus, and medium for launching satellites in emergency scenarios, comprising: if the satellite launch point is a fixed value and the orbital inclination is an unknown value, determining the first latitude argument and the geographical longitude of the first ascending node when the satellite passes through the launch point at the first current orbital inclination, and the second latitude argument and the geographical longitude of the second ascending node when passing through the observation point, based on the coordinates of the observation point and the coordinates of the launch point; determining the target orbital inclination of the satellite and the corresponding launch window based on the first latitude argument, the geographical longitude of the first ascending node, the second latitude argument, and the geographical longitude of the second ascending node; if the satellite launch point is an unknown value and the orbital inclination is a fixed value, determining the longitude and latitude of the nadir point corresponding to each launch window based on the coordinates of the observation point, the target observation time, and the observation time step; thus, in emergency launch scenarios, the launch window can be determined according to the launch requirements, thereby quickly and accurately launching the satellite to a preset position to achieve fixed-point observation of emergency events.

The invention discloses a missile launching condition multi-dimensional judgment system based on a virtual environment, which relates to the technical field of aviation man-machine work efficiency crossing and specifically comprises a data acquisition module, a twinborn prediction module, a cross-domain fusion module, a launching judgment module and a cabin visualization module. The method comprises the following steps: acquiring six-degree-of-freedom motion data of an aerial carrier and physiological state data of an operator in real time through aerial carrier equipment to form a real-time data vector; constructing an aerial carrier-missile-operator-target four-body digital twinborn model, and generating a virtual environment prediction tensor through a Monte Carlo method; splicing the real-time data vector and the virtual environment prediction tensor, performing cross-domain fusion through an attention weighting mechanism, and outputting fusion features; based on the fusion feature, outputting a comprehensive transmission permission probability, a risk dimension vector and an optimal transmission window timestamp, and establishing a transmission permission judgment mechanism; and displaying the probability cloud picture, the physiological thermodynamic diagram and the virtual ballistic trajectory in real time through cabin display equipment.

The invention relates to a space station miniaturized fast-response low-cost cargo transportation system and method, and belongs to the technical field of space station cargo transportation. The system comprises a cargo transportation aircraft, a carrier rocket and a maneuvering launching platform, the cargo transportation aircraft adopts the small carrier rocket to be launched into an orbit through land, sea and air-based launching platforms, all-phase autonomous rapid intersection and mechanical arm hovering capture butt joint are adopted, the cargo transportation aircraft enters and exits a space station through an air brake cabin, and up-down transportation of cargoes is achieved. The method has no constraint on a launching site and a launching window, and the time from task receiving to transportation to a space station does not exceed 3 days; high-precision and high-value rendezvous and docking measuring equipment and docking mechanisms do not need to be configured, high-thrust rockets are not needed either, the comprehensive cost of tasks is greatly reduced, meanwhile, the return capsule is configured, descending goods of the space station can be carried to return to the ground, platform equipment can be repeatedly used after returning, and cost is further reduced and efficiency is further improved.

The invention relates to a pre-firing and firing-direction calibration method for a firing device, and solves the problem that the process is tedious and time-consuming due to the fact that station setting, measurement and calculation usually need to be carried out for many times around a firing position in the current method. Measurement and angle comparison are carried out through a total station to ensure that a launching axis is consistent with a theoretical launching direction, and the specific process comprises the steps of setting a calibration point and measuring coordinates, erecting the total station, calculating and adjusting to the theoretical launching direction, aligning a breech and a launching window to record a horizontal angle, and finally rapidly judging the pointing deviation of the launcher by comparing angle difference values before and after lifting. Accurate measurement and angle control are emphasized in the whole process, so that the accuracy of the launching direction is ensured. According to the method, the pointing precision of the automatic gun adjusting system can be accurately evaluated, key data are provided for test result analysis and serve as the last safety insurance, and if the detected deviation exceeds an allowable threshold value, the launching process can be immediately stopped, and problems can be checked, so that serious accidents caused by system errors or faults are effectively avoided.

The application discloses a spacecraft rendezvous and docking method and device based on step-by-step orbit raising phase modulation control, and relates to the technical field of spaceflight. The spacecraft rendezvous and docking method comprises the following steps: determining the coplanar phase difference between a target spacecraft and a visiting spacecraft corresponding to each launch window based on a preset launch window range; determining a phase error reference based on a determination date; determining a target launch window from the preset launch window range based on the coplanar phase difference, the phase error reference and a target phase difference; and launching the visiting spacecraft in the target launch window based on a step-by-step orbit raising phase modulation control strategy. The application solves the technical problem that the launch window cannot be accurately determined in the spacecraft rendezvous and docking task in the related art, resulting in low phase modulation control efficiency.

The present application provides a method, a selection device and an electronic device for pre-launch selection of a passive load reduction trajectory for a spacecraft. The selection method comprises: calculating first and second high-altitude wind data at a preset altitude according to the standard high-altitude wind data used in the design of the passive load reduction trajectory and the high-altitude wind data involved in the forecast of the spacecraft launch window; calculating the comprehensive characteristic value corresponding to the standard high-altitude wind data used in the design of the passive load reduction trajectory and the comprehensive characteristic value corresponding to the high-altitude wind data of the spacecraft launch window forecast according to the average value of the first and second high-altitude wind data with altitude and the weight of the characteristic wind speed and wind direction of the spacecraft high-altitude wind; calculating the error between the comprehensive characteristic value of the standard high-altitude wind data used in the design of the passive load reduction trajectory and the comprehensive characteristic value of the high-altitude wind data of the spacecraft launch window forecast; selecting the passive load reduction trajectory scheme corresponding to the minimum error, and binding the flight parameters corresponding to the scheme to carry out the corresponding flight test mission.