30 results about "Launch window" 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 present invention belongs to the field of aerospace asteroid defense technology and relates to a full-process design method for variable-mass gravitational traction deflection of an asteroid orbit based on real-world operating constraints. The method comprises the following steps: obtaining the orbital elements corresponding to the target asteroid based on initial conditions; substituting them into a high-precision numerical integration model to calculate the perigee moment and perigee distance without gravitational traction; optimizing the launch window using a genetic algorithm and sequential quadratic programming, with the goal of maximizing the perigee deflection distance, based on warning time requirements, launch capacity, and real-world engineering constraints; and substituting the optimized launch window into a high-precision deflection efficiency numerical integration model to output the asteroid's perigee deflection distance. This method, based on the full-process mission perspective, conducts a comprehensive analysis combining warning time, launch capacity, and actual operating conditions to ensure the authenticity and engineering applicability of the calculation results.

The present invention discloses a method for designing a launch window for a low-inclination GEO satellite with the least propellant in the whole life cycle. The method comprises the following steps: obtaining input conditions; obtaining the right ascension Ω0 of the ascending node of the transfer orbit corresponding to the launch time t; obtaining the total mass of the orbit transfer propellant and the north-south orbit maintaining propellant; judging whether the total mass of the orbit transfer propellant and the north-south orbit maintaining propellant obtained in step 3 reaches the minimum, and if so, recording the corresponding total amount of propellant and the corresponding launch time t; judging the time t, if t=T f , select the launch time corresponding to the minimum total mass of the orbit transfer propellant and the north-south orbit maintenance propellant as the window front; and obtain the window rear edge according to the preset width of the launch window. The present invention can effectively improve the satellite carrying capacity and extend the service life of the satellite.

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 invention discloses a method and system for dynamically reconstructing a launch window for a lunar orbit, and relates to the technical field of lunar exploration launch orbit design. The method comprises the following steps: receiving the initial transfer orbit elements of the lunar orbit, and calculating the orbit elements at the moment of satellite-rocket separation according to the initial transfer orbit elements of the lunar orbit; inputting the orbit elements at the moment of satellite-rocket separation into an accurate dynamics equation, iteratively obtaining the accurate orbit elements of the first lunar orbit at the take-off moment, changing the take-off moment, and iteratively obtaining the accurate orbit elements of multiple lunar orbits; performing polynomial fitting on the bound elements of the accurate orbit elements of the multiple lunar orbits to obtain the polynomial coefficients corresponding to each orbit, and binding the elements of a launch vehicle as the polynomial coefficients corresponding to each orbit, so as to realize dynamic reconstruction of the lunar orbit according to the bound elements of the launch vehicle.

The present invention discloses a method for designing an orbit of a kinetic impact mission on a near-Earth asteroid with the optimal deflection effect, comprising the following steps: S1, selecting the initial parameters of the spacecraft, including the departure and arrival times of the spacecraft; S2, designing the spacecraft orbit as a splicing of three segments, namely a large elliptical Earth parking orbit, an Earth escape orbit starting from the large elliptical orbit, and an interplanetary transfer orbit, and calculating the orbit parameters; S3, evaluating the deflection effect of the target near-Earth asteroid; S4, traversing all possible departure and arrival times of the spacecraft, repeating steps S1 to S3, drawing a contour map of the deflection distance versus the departure and arrival times of the spacecraft, and determining the optimal spacecraft launch window from it after eliminating unreasonable times. By adopting a large elliptical Earth parking orbit and screening the spacecraft launch window, the present invention enables the mass and relative velocity of the spacecraft to be relatively large when impacting the asteroid, thereby achieving the optimal asteroid deflection effect.

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.

A method and device for planning a launch window and a launch point location of a carrier rocket relate to the technical field of launch of carrier rockets, and the method comprises the following steps: calculating the position of a satellite at each moment according to basic parameters of a target constellation under a J2000 coordinate system; after position information of satellites is converted into geodetic coordinates, sub-satellite point trajectories of the satellites are calculated according to the geodetic coordinates, time windows of the satellites passing through a launching site are calculated in combination with the center position and the maneuvering radius of the launching site, and key data are provided for subsequent planning of launching windows and launching point locations of a carrier rocket. The model of the carrier rocket, the time window and the sub-satellite point track of the satellite are comprehensively considered, the most suitable launching window and launching point location are obtained through iterative calculation, the tedious process of reversely deducing the launching window only by depending on the fixed launching point location and ascending node / descending node information is avoided, the modeling and calculation process is greatly simplified, and the modeling efficiency is improved. And the execution efficiency of the launching task is obviously improved.

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.

The invention relates to the technical field of spaceflight launching, and provides a rocket launching control method and system, electronic equipment and a medium. For each group of pre-injection programs, if a control instruction is abnormal or has a fault in front of a coupling point, processing the abnormality or the fault; at the coupling point corresponding to the non-last control instruction, if the control instruction is still abnormal or the fault is not eliminated, judging whether the control instruction is in the lowest emission condition range or not, and if the control instruction is not in the lowest emission condition range, abandoning the abnormal or fault part of the control instruction; and S4, when the abnormity or the fault of the control instruction exists at the last coupling point and is still not eliminated, judging whether the abnormity or the fault of the control instruction is in the lowest emission condition range or not, if not, abandoning the abnormity or the fault part of the control instruction, and igniting and emitting according to preset time. According to the scheme, the coupling point setting of various spaceflight launches can be determined, more fault and exception handling time can be gained, and the normal launch of the rocket in the launch window can be ensured.

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.

The invention provides a method and a device for selecting a passive load shedding trajectory of a spacecraft before launching, and electronic equipment, and the method comprises the steps: correspondingly calculating first and second high-altitude wind data at a preset height according to standard high-altitude wind data adopted during the design of the passive load shedding trajectory and high-altitude wind data related to the prediction of a launching window of the spacecraft; according to the average value of the first high-altitude wind data and the second high-altitude wind data along with the height and the weight of the spacecraft high-altitude wind characteristic wind speed and wind direction, the comprehensive characteristic value corresponding to the standard high-altitude wind data adopted during passive load shedding trajectory design and the comprehensive characteristic value corresponding to the spacecraft launching window forecast high-altitude wind data are calculated; calculating the error between the comprehensive characteristic value of the standard high-altitude wind data adopted during passive load shedding trajectory design and the comprehensive characteristic value of the spacecraft launching window forecast high-altitude wind data; and selecting a passive deloading trajectory scheme corresponding to the minimum error, and binding flight data corresponding to the scheme so as to carry out corresponding flight test tasks.

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 remote controller launching window assembling device, and relates to the technical field of remote controller assembling, the remote controller launching window assembling device comprises a support and a synchronous driving assembly, the synchronous driving assembly is arranged in the middle of the support, and the synchronous driving assembly comprises a back plate, double cams, a cam groove, a driven toothed bar, a driving motor, a mounting rod, a rear straight swing arm and a front V-shaped swing arm; and the back plate is fixedly mounted at the top of the rear end of the support. In the using process, on one hand, high-precision mechanical transmission of the cam is utilized, the positioning precision of the mounting positions of the two launching windows in the repeated movement is guaranteed, on the other hand, linkage design is carried out through the rotary material distributing assembly and front-back telescopic movement of a telescopic rod under transmission of the cam, and the positioning precision of the mounting positions of the two launching windows is improved. When the telescopic rod extends forwards to take materials, alternate feeding of two different types of launching windows is achieved, the vertical pitching action is overlaid when the telescopic rod contracts backwards subsequently, the assembling operation of the double-launching-window remote controller is completed, the structural design is reasonable, and the linkage performance is higher.

The utility model relates to the field of simulation practice equipment, in particular to an equipment simulation practice cabin which comprises a cabin body, the cabin body is provided with an access door and a launching window, the inner side wall of the cabin body simulates the inner side wall structure of equipment, an operation table is arranged in the cabin body, and a simulation launching mechanism is arranged on the operation table. A plurality of simulation equipment buckles are arranged on the inner wall of the cabin body, and the simulation equipment buckles are used for placing practice equipment which simulates the launching mechanism to launch. According to the utility model, the structure is reasonable, the main real environment inside real equipment can be simulated, the practice equipment is placed on the simulation equipment buckle, and during practice, the practice equipment is taken down from the simulation equipment buckle and placed on the simulation launching mechanism for practice filling, so that the filling speed can be increased through repeated practice, and the practice effect is achieved; and the problem of damage caused by frequent practice by using real equipment is avoided.

The invention discloses a launch window dynamic reconstruction design method and system for a moon-striving orbit, and relates to the technical field of lunar exploration launch orbit design, and the method comprises the following steps: receiving the number of initial transfer orbits of the moon-striving orbit, and calculating the number of orbits at a satellite-rocket separation point moment according to the number of the initial transfer orbits of the moon-striving orbit; inputting the orbital element at the satellite-rocket separation point moment into the precise kinetic equation, carrying out iteration to obtain the precise orbital element of the first moon-striving orbit at the takeoff moment, changing the takeoff moment, and carrying out iteration to obtain the precise orbital elements of a plurality of moon-striving orbits; and performing polynomial fitting on the accurate orbit number binding data of the multiple moon-striving orbits to obtain a polynomial coefficient corresponding to each orbit, and binding data of the carrier rocket by taking the polynomial coefficient as the polynomial coefficient corresponding to each orbit, thereby realizing dynamic reconstruction window moon-striving according to the binding data of the carrier rocket.

Embodiments of the present application provide an application launching method and an electronic device. The method is applied to an electronic device comprising a system service process, a system interface process, and a desktop process, and the method comprises: in response to a first operation of a user, the desktop process sends a launch message to the system service process, the launch message indicating a first function of launching a first application by means of a shortcut, and an interface corresponding to the first function being an interface displayed after the first application sequentially loads n activities; in response to the launch message, the system service process launches a first activity among the n activities, and creates a first transition component, wherein in the process of launching the first activity, the system interface process adds and draws a first launch window for the first activity; in response to the first transition component being ready, the system interface process notifies the desktop process to execute the first transition component; and the desktop process executes the first transition component. The method can increase the response speed during application launching.

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 the field of spaceflight launch carrier rockets, in particular to a pre-launch temperature prediction calculation method and system for a carrier rocket component, and the method comprises the steps: building a heat conduction simulation model of the carrier rocket component; carrying out an environment test on the carrier rocket component in a laboratory, and obtaining a thermal diffusion coefficient of the heat conduction simulation model through the environment test; according to the obtained thermal diffusion coefficient of the heat conduction simulation model, establishing a relation model among the temperature change of the carrier rocket component, the temperature of a carrier rocket storage tank propellant and the environment temperature; and inputting the environment temperature monitored before launching and the propellant temperature into the relation model so as to predict and calculate the temperature of the carrier rocket component at any moment. The temperature and the change trend of the carrier rocket component can be quickly predicted and calculated, whether the carrier rocket component meets the temperature condition of the launching window or not is determined, and a decision basis is provided for a commander.

The present invention discloses a method for planning the rendezvous launch window of a near-Earth spacecraft, which can analyze the launch windows of the tracking and target spacecraft when neither the target spacecraft nor the tracking spacecraft has been launched. First, corresponding to the initial epoch moment, an objective orbit with the initial right ascension of the ascending node corresponding to the initial epoch moment is established. A variation curve of the angle between the orbital plane and the illumination position vector within a period of time is constructed; according to the illumination conditions and the time period during the launch of the on-orbit spacecraft, the time periods that meet the illumination constraints are screened out. The illumination position vector is the position vector from the center of mass of the on-orbit spacecraft to the center of mass of the sun. According to the launch site positions of the target spacecraft and the tracking spacecraft, the launch descending orbit zero windows of the tracking spacecraft and the target spacecraft are found, which are the launch window lists corresponding to the initial right ascension of the ascending node at the initial epoch moment. Traverse the initial right ascension of the ascending node. For different initial right ascensions of the ascending node, launch window lists corresponding to different initial right ascensions of the ascending node are obtained.

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 invention discloses a design method for the elements of achieving takeoff at any time during the Mars exploration launch window, which relates to the technical field of Mars exploration launch orbit design. The present invention includes: calculating the initial values required for the departure parameters of the detector according to the launch time; designing the orbital parameters of the launch vehicle according to the initial values of the detector departure parameters, and combining the calculation amount and design amount in the orbital parameters of the launch vehicle to complete the launch orbit design through iteration; adjusting the target orbit inclination and calculating and generating the target elements of the launch orbit with strong regularity of change with the takeoff time. The present invention can enable the launch vehicle to achieve takeoff at any time within the launch window on the same day, meet the precise departure parameter requirements for the detector to escape and fly on the orbit to Mars, save the additional propellant consumption of the detector caused by the deviation of the takeoff time after the detector separation, improve the transportation capacity of the detector platform, and at the same time improve the design work 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.