Method for correcting wind before spaceflight launch vehicle

Abstract

The invention provides a method for correcting wind before launching of an aerospace vehicle. The method comprises the following steps: acquiring high-altitude wind data actually measured in minus 1 hour; taking minus 1 hour high-altitude wind data as input of trajectory correction, and outputting guidance parameters; carrying out six-degree-of-freedom Monte Carlo shooting simulation, and calculating a flight load; and determining whether to launch the space launch vehicle at least according to the flight load envelope. The method can improve the launching release probability and reduce the pneumatic attack angle and the flight load.

Description

Technical Field

[0001] The present invention relates to the technical field of aerospace engineering, and in particular to a method for wind correction before launch of a space vehicle. Background Art

[0002] In the existing technology, the wind correction technology for space launch vehicles is based on the statistical high-altitude wind field to design multiple wind-corrected trajectories. That is, multiple trajectories are designed in advance based on historical statistical wind fields during the design phase, such as 0%, 25%, 50%, and 75% wind speed correction levels. On the launch day, the optimal solution is selected from the pre-designed trajectories based on the actual measured wind data before launch.

[0003] Statistical wind field modeling, based on historical data averages or extreme values, cannot accurately reflect the transient wind shear, turbulence, and other characteristics of the real-time wind field on launch day. This results in non-zero actual flight angles of attack and conservative payload calculations. Pre-designed multi-stage trajectories cannot fully match the measured wind field on launch day. For example, the traditional four-stage wind correction technology fails to meet launch conditions in approximately 5% of historical wind field scenarios.

[0004] In view of this, there is an urgent need to design a method for pre-launch wind correction of space launch vehicles that is based on the measured high-altitude wind field and can accurately calculate the payload. Summary of the Invention

[0005] The purpose of the present invention is to overcome the deficiencies of the prior art and provide a method for wind correction before launch of a space vehicle.

[0006] The present invention provides a method for wind correction before a space vehicle is launched, comprising: collecting negative one-hour measured high-altitude wind data; using the negative one-hour high-altitude wind data as input for trajectory correction and outputting guidance parameters; performing six-degree-of-freedom Monte Carlo target shooting simulation and calculating flight load; and determining whether to launch the space vehicle based at least on the flight load envelope.

[0007] According to one embodiment of the present invention, before the step of collecting the actual measured high-altitude wind data for -1 hour, the step also includes: collecting the actual measured high-altitude wind data for -3 hours; using the actual measured high-altitude wind data for -3 hours as the input for trajectory correction and outputting guidance parameters; conducting a six-degree-of-freedom Monte Carlo target shooting simulation and calculating the flight load; and deciding whether to refuel the space vehicle based at least on the flight load envelope.

[0008] According to one embodiment of the present invention, the collection of high-altitude wind data measured in the negative 1 hour and the collection of high-altitude wind data measured in the negative 3 hours at least includes: collecting the distribution data of wind speed and wind direction measured in the launch area with altitude.

[0009] According to one embodiment of the present invention, the method of using negative 1-hour high-altitude wind data as the input of ballistic correction and using negative 3-hour high-altitude wind data as the input of ballistic correction is specifically as follows: solving the ballistic motion equation based on the measured high-altitude wind data with the minimum aerodynamic angle of attack as the target, wherein the longitudinal axis of the space vehicle is made consistent with the airspeed vector in the transonic section and the maximum dynamic pressure area, while satisfying the orbital entry condition constraints.

[0010] According to one embodiment of the present invention, the output guidance parameters are specifically: based on the quasi-real-time wind-modified trajectory, at least the guidance parameters related to the program angle, perturbation guidance coefficient, standard guidance amount, iterative guidance parameters and shutdown equation coefficients are generated.

[0011] According to one embodiment of the present invention, the six-degree-of-freedom Monte Carlo target shooting simulation method is specifically as follows: random errors of at least inertial group noise and wind field deviation are injected in a normal distribution manner, and combined navigation corrections are made using simulated satellite navigation simulation to simulate real flight conditions to obtain flight area safety, landing area safety, propellant remainder and orbit insertion accuracy.

[0012] According to one embodiment of the present invention, the method for calculating the flight load is specifically: extracting the data of the aerodynamic angle of attack and the engine swing angle in the flight phase in the six-degree-of-freedom Monte Carlo target simulation method, calculating the lateral normal aerodynamic load, axial static load, sway load, gust load and buffeting load, and comprehensively obtaining the equivalent axial pressure and axial tension load envelopes of each section of the rocket body.

[0013] According to one embodiment of the present invention, in addition to the flight payload envelope, flight area safety, landing area safety, propellant remainder and orbit insertion accuracy all need to meet launch requirements.

[0014] According to one embodiment of the present invention, the safety of the navigation area means that the flight trajectory of the six-degree-of-freedom Monte Carlo shooting simulation does not exceed the safety pipe boundary; the safety of the landing area means that the first sub-stage and fairing wreckage in the six-degree-of-freedom Monte Carlo shooting simulation are located within the landing area.

[0015] According to one embodiment of the present invention, the propellant remainder is such that the probability of a sub-stage being depleted is no more than 15%, and the probability of a second-stage being depleted before entering orbit is 0; the orbit insertion accuracy is such that the deviation of the orbit semi-major axis is no more than 5 km, the deviation of the orbit eccentricity is no more than 0.002, and the deviation of the orbit inclination is no more than 0.08°.

[0016] In this embodiment, the measured high-altitude wind data of -1 hour is selected as the launch window wind field, and the high-altitude wind correction trajectory design is carried out based on this. The payload envelope is closer to the actual flight, avoiding the "margin stacking" of the traditional method, reducing the aerodynamic angle of attack and flight load, improving the launch release probability, and expanding the adaptability of the launch window.

[0017] It should be understood that the foregoing general description and the following detailed description are merely exemplary and illustrative and are not intended to limit the scope of the present invention. BRIEF DESCRIPTION OF THE DRAWINGS

[0018] The following drawings are a part of the specification of the present invention and illustrate exemplary embodiments of the present invention. Together with the description, the accompanying drawings serve to explain the principles of the invention.

[0019] Figure 1 Schematic diagram of a method for pre-launch wind correction of a space vehicle according to an embodiment of the present invention; Figure 2 is a schematic diagram of a method for pre-launch wind correction of a space vehicle according to another embodiment of the present invention; Figure 3 This is a comparison diagram between the measured wind field and the statistical wind field in one embodiment of the present invention. DETAILED DESCRIPTION

[0020] The features and exemplary embodiments of various aspects of the present invention will be described in detail below. In order to make the purposes, technical solutions and advantages of the present invention more clear, the present invention will be further described in detail below in conjunction with the accompanying drawings and specific embodiments. It should be understood that the specific embodiments described herein are only configured to explain the present invention and are used to illustrate the principles of the present invention, and are not configured to limit the present invention. In addition, the structural components in the drawings are not necessarily drawn to scale. For example, the dimensions of some structural components or areas in the drawings may be enlarged for other structural components or areas to facilitate understanding of the embodiments of the present invention.

[0021] The directional words appearing in the following description refer to the directions shown in the drawings and do not limit the specific structure of the embodiments of the present invention. In the description of the present invention, it should be noted that, unless otherwise specified, the terms "installed", "connected", and "connected" should be understood in a broad sense. For example, it can be a fixed connection, a detachable connection, or an integral connection; it can be directly connected or indirectly connected through an intermediate medium. For those skilled in the art, the specific meanings of the above terms in the present invention can be understood according to the specific circumstances.

[0022] In addition, the terms "include", "comprising", "having" or any other variations thereof are intended to cover non-exclusive inclusion, so that a structure or component comprising a series of elements includes not only those elements, but also other mechanical elements not explicitly listed or inherent in the structure or component. In the absence of more limitations, the elements defined by the sentence "comprising..." do not exclude the presence of other identical elements in the article or device comprising the elements.

[0023] Spatially relative terms such as "below," "beneath," "under," "low," "above," "on," "high," and the like are used to facilitate description to explain the positioning of one element relative to a second element, indicating that these terms are intended to encompass different orientations of the device in addition to orientations different from those shown in the figures. In addition, for example, "one element is above / below another element" can mean that the two elements are in direct contact or that there are other elements between the two elements. Furthermore, terms such as "first," "second," and the like are also used to describe various elements, regions, portions, and the like, and should not be considered limiting. Similar terms are used throughout the description to represent similar elements.

[0024] For those skilled in the art, the present invention can be implemented without some of these specific details. The following description of the embodiments is only intended to provide a better understanding of the present invention by showing examples of the present invention.

[0025] Figure 1 Schematic diagram of a method for pre-launch wind correction of a space vehicle according to an embodiment of the present invention; Figure 2 is a schematic diagram of a method for pre-launch wind correction of a space vehicle according to another embodiment of the present invention; Figure 3 This is a comparison diagram between the measured wind field and the statistical wind field in one embodiment of the present invention.

[0026] like Figure 1 As shown, the present invention provides a method for wind correction before launch of a space vehicle, comprising: Step S101, collecting negative 1 hour measured upper air wind data; Step S102, using the negative 1-hour upper-altitude wind data as the input for trajectory correction and outputting guidance parameters; Step S103, performing six-degree-of-freedom Monte Carlo target shooting simulation to calculate flight load; Step S104: determining whether to launch the space vehicle based at least on the flight payload envelope.

[0027] like Figure 2 As shown, according to one embodiment of the present invention, before the step of collecting the negative 1 hour measured high-altitude wind data, the method further includes: Step S201, collecting negative 3-hour measured upper-altitude wind data; Step S202, using the negative 3-hour upper-altitude wind data as the input for trajectory correction and outputting guidance parameters; Step S203, performing a six-degree-of-freedom Monte Carlo target shooting simulation to calculate the flight load; Step S204 , determining whether to refuel the space vehicle based at least on the flight payload envelope.

[0028] Among them, in the traditional wind correction technology, the problems of large flight load of space vehicle and low release probability on launch day due to mismatch between statistical wind field and measured wind field and conservative multi-speed trajectory design can be improved in this embodiment. Figure 3 As shown, the diamond dotted line in the statistical wind field model is the statistical maximum, the square dotted line is 75% of the statistical maximum, the triangle dotted line is 50% of the statistical maximum, the circular dotted line is 25% of the statistical maximum, and the solid line is the measured wind speed value. It can be seen that the wind direction and wind speed of the measured high-altitude wind field are very different from those of the statistical wind field model. The measured high-altitude wind field can accurately reflect the complex wind field characteristics such as transient wind shear and turbulence in actual flight, significantly improve the characterization accuracy of wind shear gradient and transient turbulence characteristics, and overcome the inherent limitations of traditional statistical wind field models in simulating complex meteorological conditions. The method in this embodiment is based on quasi-real-time ballistic correction of the measured wind field, and the load envelope is closer to actual flight, avoiding the "margin superposition" of the traditional method, reducing the aerodynamic angle of attack and flight load, improving the launch release probability, and expanding the adaptability of the launch window.

[0029] Unlike the traditional four-speed wind correction technology, the trajectory design in the quasi-real-time wind correction technology does not use statistical wind profiles, but directly uses the measured wind before launch as the trajectory design input. Quasi-real-time wind correction technology can replace the previous statistical wind field, greatly improving the accuracy of wind field representation and reducing the error in angle of attack calculation from the source. When a space vehicle flies in a strong wind area according to the quasi-real-time wind correction trajectory, the theoretical angle of attack is close to 0°, which fundamentally reduces the flight load of the space vehicle. Under ideal conditions, when the input high-altitude wind field data is the wind field when the space vehicle actually flies through a strong wind area, the aerodynamic angle of attack is 0°.

[0030] In practice, it is impossible to obtain the wind speed vector of a space vehicle in real time during flight and design the corresponding wind-corrected trajectory. In this embodiment, the measured high-altitude wind data for a period of time before the launch window is selected. For example, in this embodiment, the measured high-altitude wind data for -1 hour is selected as the launch window wind field, and the high-altitude wind-corrected trajectory design is carried out based on this, that is, the quasi-real-time wind-corrected trajectory design. It can be seen that the closer the time to the launch window is when the high-altitude wind data is obtained, the better the wind correction effect. Therefore, in the implementation process, the reasonable selection of the time to measure the high-altitude wind field is an important link to ensure the quasi-real-time wind correction effect.

[0031] Optionally, this embodiment may use machine learning to assist in trajectory design, using trajectory data from historical wind fields to train a neural network to achieve rapid optimization of trajectory correction parameters, further shortening the design time.

[0032] Specifically, the launch vehicle visits the meteorological office to obtain negative three-hour upper-altitude wind data. This data is then transmitted back as input for trajectory corrections. Based on this negative three-hour upper-altitude wind data, near-real-time wind correction design, simulation, analysis, and judgment are performed. Subsequently, flight payloads are calculated through six-degree-of-freedom Monte Carlo target simulation, resulting in a comprehensive load envelope for each section of the space vehicle. The propellant loading requirements for the space vehicle are determined based on at least the value of this load envelope. If the loading conditions are met, guidance data and a notification form are output. If the loading conditions are not met, the launch process is aborted.

[0033] Next, the space vehicle launcher goes to the meteorological room to obtain the negative 1-hour high-altitude wind data, and transmits the high-altitude wind data back as the input for trajectory correction. Based on the negative 1-hour high-altitude wind data, quasi-real-time wind correction design, simulation, analysis and judgment are carried out. The flight load is calculated after the six-degree-of-freedom Monte Carlo target shooting simulation, and the load envelope of each section of the space vehicle is comprehensively obtained. The release of the space vehicle is determined at least based on the value of the load envelope. If the release criteria are met, the quasi-real-time wind-corrected trajectory and guidance data are output. After entering the launch process, the negative 1-hour launch data are transmitted to the space vehicle control command post. After inspection, format conversion, uploading and other steps, the guidance data binding is finally completed. If the launch conditions are not met, the launch process is terminated.

[0034] The entire process in this embodiment has a rapid response mechanism. From the acquisition of wind field data to the binding of guidance data, the entire process takes less than half an hour, meeting the timeliness requirements of launch day decision-making and supporting quasi-real-time trajectory adjustment.

[0035] According to one embodiment of the present invention, collecting the high-altitude wind data measured in the negative 1 hour and collecting the high-altitude wind data measured in the negative 3 hours at least includes: collecting the distribution data of the wind speed and wind direction measured in the launch area with altitude.

[0036] According to one embodiment of the present invention, a method of using negative 1-hour high-altitude wind data as input for trajectory correction and a method of using negative 3-hour high-altitude wind data as input for trajectory correction is specifically as follows: solving the ballistic motion equation based on the measured high-altitude wind data with the minimum aerodynamic angle of attack as the target, wherein the longitudinal axis of the space vehicle is made consistent with the airspeed vector in the transonic section and the maximum dynamic pressure area, while satisfying the orbital entry condition constraints.

[0037] According to one embodiment of the present invention, the output guidance parameters are specifically: based on the quasi-real-time wind-modified trajectory, at least the guidance parameters related to the program angle, perturbation guidance coefficient, standard guidance amount, iterative guidance parameters and shutdown equation coefficients are generated.

[0038] According to one embodiment of the present invention, a six-degree-of-freedom Monte Carlo target shooting simulation method is specifically as follows: random errors of at least inertial group noise and wind field deviation are injected in a normal distribution manner, and combined navigation corrections are made using simulated satellite navigation simulation to simulate real flight conditions to obtain flight area safety, landing area safety, propellant remainder, and orbit insertion accuracy.

[0039] According to one embodiment of the present invention, a method for calculating flight loads is specifically as follows: extracting data on aerodynamic attack angles and engine swing angles during the flight phase in a six-degree-of-freedom Monte Carlo target practice simulation method, calculating transverse normal aerodynamic loads, axial static loads, sway loads, gust loads, and buffeting loads, and comprehensively obtaining the equivalent axial compressive and axial tensile load envelopes of each section of the rocket body.

[0040] According to one embodiment of the present invention, in addition to the flight payload envelope, flight area safety, landing area safety, propellant remainder and orbit insertion accuracy all need to meet launch requirements.

[0041] According to one embodiment of the present invention, the safety of the flight area means that the flight trajectory of the six-degree-of-freedom Monte Carlo shooting simulation does not exceed the safety pipeline boundary; the safety of the landing area means that the first sub-stage and fairing wreckage in the six-degree-of-freedom Monte Carlo shooting simulation are located within the landing area.

[0042] According to one embodiment of the present invention, the propellant remainder is such that the probability of a sub-stage being depleted is no more than 15%, and the probability of a second-stage being depleted before entering orbit is 0; the orbit insertion accuracy is such that the deviation of the orbit semi-major axis is no more than 5 km, the deviation of the orbit eccentricity is no more than 0.002, and the deviation of the orbit inclination is no more than 0.08°.

[0043] Specifically, based on the measured wind field's wind speed and direction distribution data with altitude, the ballistic motion equations are solved with the minimum aerodynamic angle of attack as the goal, and a wind-corrected trajectory is designed. This wind-corrected trajectory requires that the longitudinal axis of the spacecraft be aligned with the airspeed vector in the transonic region and the maximum dynamic pressure zone, meaning that the theoretical angle of attack approaches 0°, while also satisfying orbital insertion constraints. After obtaining a quasi-real-time wind-corrected trajectory scheme based on the measured pre-launch wind field, the guidance parameters are calculated, including the characteristic quantities of the program angle, perturbation guidance coefficient, standard guidance amount, iterative guidance parameters, and shutdown equation coefficients. All of these guidance parameters are automatically designed.

[0044] Next, in order to verify whether the quasi-real-time wind-corrected trajectory scheme and the guidance parameters and various parameters designed based on the quasi-real-time wind-corrected trajectory meet the mission requirements, a six-degree-of-freedom Monte Carlo target shooting simulation verification is carried out. The target shooting simulation test randomly injects various errors according to the normal distribution method, introduces the inertial group noise and inertial group model, and uses simulated satellite guidance to simulate the combined navigation correction. For example, in one embodiment, more than 2,000 six-degree-of-freedom simulations can be carried out to simulate real flight conditions. Considering the impact of the landing area wind field on the landing point position of the first stage and the fairing, the landing area is used to calculate the high-altitude wind field in the design phase. When performing the mission, the landing area prediction wind field of the launch day minus 1 day is used. Through simulation verification, the guidance system correctly outputs instructions, the attitude control system manipulates the servo mechanism, swings the engine, and controls the space vehicle to fly stably according to the quasi-real-time wind-corrected trajectory.

[0045] Flight load calculations primarily focus on the phase when the spacecraft transits high winds, i.e., 30 to 110 seconds after takeoff. Data such as aerodynamic angle of attack and engine swing angle are obtained from six-degree-of-freedom Monte Carlo target simulations. Transverse normal aerodynamic loads, axial static loads, sway loads, gust loads, and buffeting loads are calculated, comprehensively deriving the equivalent axial compressive and tensile load envelopes for each section of the vehicle. Axial static loads are determined by engine thrust, aerodynamic drag, and longitudinal overload. Transverse normal aerodynamic loads must account for sway loads caused by the liquid in the tank, elastic response loads caused by gust disturbances (i.e., gust loads), and fluctuating pressure loads (i.e., buffeting loads) in the transonic region. By superimposing static and dynamic loads at the same time, axial forces, shear forces, and bending moments are determined for each section of the spacecraft, from which equivalent axial compressive and tensile forces are calculated. Load assessments based on Monte Carlo simulation results quantify load distribution characteristics and avoid conservative design.

[0046] Optionally, this embodiment may adopt a multi-objective joint optimization technology to couple payload reduction with indicators such as orbit insertion accuracy and propellant consumption, and achieve comprehensive performance improvement through optimization tools such as genetic algorithms.

[0047] In specific operations, in order to apply quasi-real-time wind correction technology to determine whether the high-altitude wind field meets the launch requirements before launch insolation, corresponding release criteria need to be formulated. The release criteria are composed of parameters and criteria, and mainly consider flight area safety, landing area safety, propellant remaining, orbit insertion accuracy, and flight load factors. Among them, flight area safety requires that the six-degree-of-freedom Monte Carlo target shooting simulation flight trajectory does not exceed the safety pipeline boundary; landing area safety requires that the first stage and fairing debris are located within the landing area; the propellant safety margin requires that the probability of the first stage being shut down is no more than 15%, and the probability of the second stage being shut down before entering orbit is 0; the satellite orbit insertion accuracy requires that the orbit semi-major axis deviation is no more than 5km, the orbit eccentricity deviation is no more than 0.002, and the orbit inclination deviation is no more than 0.08°.

[0048] It should be noted that prior to the launch mission, the allowable load limits for each structural segment of the spacecraft were determined based on the ultimate load-bearing capacity, safety factor, residual strength coefficient, and launch reliability requirements for each segment of the spacecraft's structural products during flight in Category 1 winds. Based on the measured pre-launch winds, near-real-time wind-correction design and simulation were conducted to predict the equivalent axial compressive and axial tensile load envelopes for each structural segment during flight in Category 1 winds. When the predicted pre-launch load envelope is less than the allowable load limit, the load release conditions are met.

[0049] Adopting the method in this embodiment can significantly increase the launch release probability of a space vehicle, from 95% with traditional technology to 99.95%, reducing launch delay costs. This method integrates multidisciplinary models for wind field, trajectory, payload, and control, enabling closed-loop optimization from wind field measurement to launch decision-making, improving system design efficiency. This method supports launches under more stringent wind field conditions, enhances space vehicle mission flexibility, and adapts to the high-frequency launch requirements of commercial spaceflight. In this embodiment, the reduced payload achieved allows the design weight of the rocket body structure to be reduced, thereby improving the overall load factor.

[0050] The above are only preferred embodiments of the present invention and are not intended to limit the present invention. Any modifications, equivalent replacements, improvements, etc. made within the spirit and principles of the present invention should be included in the scope of protection of the present invention.

Claims

1. A method for wind correction before launch of a space vehicle, characterized in that: include: Collect negative 1 hour measured upper air wind data; The negative 1-hour upper-air wind data is used as the input for trajectory correction and the guidance parameters are output; Conduct six-degree-of-freedom Monte Carlo target shooting simulations to calculate flight loads; The decision to launch a space vehicle is based at least on the flight payload envelope.

2. The method for wind correction before launch of a space vehicle according to claim 1, characterized in that: Before collecting the negative 1 hour measured upper air wind data, the following steps are also included: Collect negative 3-hour measured upper-air wind data; The negative 3-hour upper-air wind data is used as the input for trajectory correction and the guidance parameters are output; Conduct six-degree-of-freedom Monte Carlo target shooting simulations to calculate flight loads; The decision to refuel the space vehicle is based at least on the flight payload envelope.

3. The method for wind correction before launch of a space vehicle according to claim 2, characterized in that: The high-altitude wind data collected in the negative 1 hour actual high-altitude wind data and the high-altitude wind data collected in the negative 3 hours actual high-altitude wind data at least include: collecting the distribution data of the wind speed and wind direction measured in the launch area with height.

4. The method for wind correction before launch of a space vehicle according to claim 3, characterized in that: The method of using the negative 1 hour high-altitude wind data as the input of the trajectory correction and the method of using the negative 3 hours high-altitude wind data as the input of the trajectory correction are specifically as follows: Based on the measured high-altitude wind data, the ballistic motion equation is solved with the minimum aerodynamic angle of attack as the goal. In the transonic section and the maximum dynamic pressure area, the longitudinal axis of the space vehicle is made consistent with the airspeed vector, and the orbital insertion conditions are met at the same time.

5. The method for wind correction before launch of a space vehicle according to claim 4, characterized in that: The output guidance parameters are specifically: According to the quasi-real-time wind-modified trajectory, at least guidance parameters related to the program angle, perturbation guidance coefficient, standard guidance amount, iterative guidance parameter and shutdown equation coefficient are generated.

6. The method for wind correction before launch of a space vehicle according to claim 5, characterized in that: The six-degree-of-freedom Monte Carlo target practice simulation method is specifically as follows: Random errors such as at least inertial group noise and wind field deviation are injected according to the normal distribution method, and the combined navigation correction is simulated by simulated satellite navigation to simulate the actual flight situation, so as to obtain the flight area safety, landing area safety, propellant remainder and orbit insertion accuracy.

7. The method for wind correction before launch of a space vehicle according to claim 6, characterized in that: The method for calculating the flight load is specifically as follows: The data of aerodynamic angle of attack and engine swing angle in the flight phase of the six-degree-of-freedom Monte Carlo target simulation method are extracted, and the lateral normal aerodynamic load, axial static load, sway load, gust load and buffeting load are calculated to comprehensively obtain the equivalent axial compression and axial tension load envelopes of each section of the rocket body.

8. The method for wind correction before launch of a space vehicle according to claim 6, characterized in that: In addition to the flight payload envelope, flight area safety, landing area safety, propellant remainder and orbit insertion accuracy must all meet launch requirements.

9. The method for wind correction before launch of a space vehicle according to claim 8, characterized in that: The safety of the navigation area means that the flight trajectory of the six-degree-of-freedom Monte Carlo shooting simulation does not exceed the safety pipeline boundary; the safety of the landing area means that the first stage and fairing wreckage in the six-degree-of-freedom Monte Carlo shooting simulation are located within the landing area.

10. The method for wind correction before launch of a space vehicle according to claim 8, characterized in that: The propellant surplus is such that the probability of a first stage being depleted is no more than 15%, and the probability of a second stage being depleted before entering orbit is 0; the orbit insertion accuracy is such that the deviation of the orbit semi-major axis is no more than 5 km, the deviation of the orbit eccentricity is no more than 0.002, and the deviation of the orbit inclination is no more than 0.08°.

Images