A method for pre-launch wind correction of a space launch vehicle

Abstract

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

Description

Technical Field

[0001] This invention relates to the field of aerospace engineering technology, and specifically to a method for pre-launch wind correction of a space launch vehicle. Background Technology

[0002] In existing technologies, the wind correction technology for space launch vehicles is based on statistical upper-altitude wind fields 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 wind speed correction levels of 0%, 25%, 50%, and 75%. On the launch day, the optimal scheme is selected from the pre-designed trajectories based on the pre-launch measured wind data.

[0003] Statistical wind field modeling, based on historical averages or extreme values, cannot accurately reflect the transient wind shear, turbulence, and other characteristics of the real-time wind field on the launch day. This results in a non-zero actual flight angle of attack and a conservative load calculation. The pre-designed multi-stage trajectory cannot fully match the measured wind field on the launch day. For example, the traditional four-stage wind correction technique cannot meet the release conditions in about 5% of historical wind field scenarios.

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

[0005] The purpose of this invention is to overcome the shortcomings of the prior art and provide a method for pre-launch wind correction of a space launch vehicle.

[0006] This invention provides a method for pre-launch wind correction of a space launch vehicle, comprising: collecting measured high-altitude wind data for -1 hour; using the -1 hour high-altitude wind data as input for trajectory correction and outputting guidance parameters; conducting a six-degree-of-freedom Monte Carlo target simulation to calculate the flight load; and determining whether to launch the space launch vehicle based at least on the flight load envelope.

[0007] According to one embodiment of the present invention, before the step of collecting measured upper-air wind data for -1 hour, the method further includes: collecting measured upper-air wind data for -3 hours; using the upper-air wind data for -3 hours as input for ballistic correction and outputting guidance parameters; conducting a six-degree-of-freedom Monte Carlo target simulation to calculate the flight load; and determining whether to refuel the space launch vehicle based at least on the flight load envelope.

[0008] According to one embodiment of the present invention, the high-altitude wind data collected during the negative 1 hour and the negative 3 hour periods include at least the following: the distribution data of wind speed and wind direction with altitude in the launch area.

[0009] According to an embodiment of the present invention, the method of using negative 1 hour upper-altitude wind data as input for ballistic correction and negative 3 hour upper-altitude wind data as input for ballistic correction specifically involves: solving the ballistic motion equation based on measured upper-altitude wind data with the minimum aerodynamic angle of attack as the objective, wherein the longitudinal axis of the space launch vehicle is aligned with the airspeed vector in the transonic band and the maximum dynamic pressure region, while simultaneously satisfying the orbital insertion condition constraints.

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

[0011] According to one embodiment of the present invention, the method of six-degree-of-freedom Monte Carlo target simulation specifically involves: injecting random errors, including at least inertial group noise and wind field deviation, in a normal distribution manner, and correcting them using simulated satellite navigation combined navigation to simulate real flight conditions, thereby obtaining flight area safety, landing area safety, propellant reserve, and orbital insertion accuracy.

[0012] According to an embodiment of the present invention, the method for calculating flight loads specifically involves: extracting data on aerodynamic angle of attack and engine sway angle during the flight phase in the six-degree-of-freedom Monte Carlo shooting simulation method; calculating lateral normal aerodynamic loads, axial static loads, sway loads, gust loads, and flutter loads; and comprehensively obtaining the equivalent axial compression and axial tension load envelopes for each section of the rocket body.

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

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

[0015] According to one embodiment of the present invention, the propellant reserve is such that the probability of first-stage exhaustion is no greater than 15%, and the probability of second-stage exhaustion before orbit insertion is 0; the orbit insertion accuracy is such that the deviation of the semi-major axis of the orbit is no greater than 5 km, the deviation of the orbital eccentricity is no greater than 0.002, and the deviation of the orbital inclination angle is no greater than 0.08°.

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

[0017] It should be understood that the above general description and the following specific embodiments are merely exemplary and illustrative, and do not limit the scope of the invention. Attached Figure Description

[0018] The accompanying drawings, which are part of the specification of this invention, illustrate exemplary embodiments of the invention. The drawings, together with the description in the specification, serve to illustrate the principles of the invention.

[0019] Figure 1 This is a schematic diagram of a method for pre-launch wind correction of a space launch vehicle according to an embodiment of the present invention;

[0020] Figure 2 This is a schematic diagram of a pre-launch wind correction method for a space launch vehicle according to another embodiment of the present invention;

[0021] Figure 3 This is a comparison chart of measured wind field and statistical wind field in one embodiment of the present invention. Detailed Implementation

[0022] The features and exemplary embodiments of various aspects of the present invention will now be described in detail. To make the objectives, technical solutions, and advantages of the present invention clearer, the present invention will be further described in detail below with reference to 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 to exemplify 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 regions in the drawings may be enlarged for other structural components or regions to aid in the understanding of the embodiments of the present invention.

[0023] The directional terms used in the following description refer to the directions shown in the figures and are not intended to 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 stated, the terms "installation," "connection," and "joining" should be interpreted broadly. For example, they can refer to a fixed connection, a detachable connection, or an integral connection; they can refer to a direct connection or an indirect connection through an intermediate medium. Those skilled in the art can understand the specific meaning of the above terms in the present invention according to the specific circumstances.

[0024] Furthermore, the terms "comprising," "including," "having," or any other variations thereof are intended to cover non-exclusive inclusion, such that a structure or component that includes a list of elements includes not only those elements but also other structural elements that are not expressly listed or inherent to the structure or component. Without further limitations, an element defined by the phrase "comprising..." does not exclude the presence of other identical elements in the article or apparatus that includes the element.

[0025] Spatial relation terms such as "below," "under," "under," "low," "above," "on," and "high" are used for descriptive convenience to explain the positioning of one element relative to a second element, indicating that these terms are intended to cover different orientations of the device, in addition to those different from those shown in the figure. Furthermore, phrases such as "one element on / below another element" can indicate that two elements are in direct contact, or that there are other elements between the two elements. In addition, terms such as "first" and "second" are also used to describe individual elements, areas, parts, etc., and should not be considered limiting. Similar terms are used throughout the description to refer to similar elements.

[0026] It will be apparent to those skilled in the art that the present invention can be practiced without requiring some of these specific details. The following description of embodiments is merely intended to provide a better understanding of the invention by illustrating examples of the invention.

[0027] Figure 1 This is a schematic diagram of a method for pre-launch wind correction of a space launch vehicle according to an embodiment of the present invention; Figure 2 This is a schematic diagram of a pre-launch wind correction method for a space launch vehicle according to another embodiment of the present invention; Figure 3 This is a comparison chart of measured wind field and statistical wind field in one embodiment of the present invention.

[0028] like Figure 1 As shown, the present invention provides a method for pre-launch wind correction of a space launch vehicle, comprising:

[0029] Step S101: Collect measured upper-level wind data for -1 hour;

[0030] Step S102: Use the negative 1 hour upper-level wind data as input for ballistic correction and output guidance parameters;

[0031] Step S103: Conduct a six-degree-of-freedom Monte Carlo target simulation and calculate the flight load;

[0032] Step S104: Determine whether to launch the spacecraft based at least on the flight payload envelope.

[0033] like Figure 2 As shown, according to one embodiment of the present invention, the method further includes the following steps before collecting negative 1 hour measured upper-level wind data:

[0034] Step S201: Collect measured upper-level wind data for -3 hours;

[0035] Step S202: Use the negative 3-hour high-altitude wind data as input for ballistic correction and output guidance parameters;

[0036] Step S203: Conduct a six-degree-of-freedom Monte Carlo target simulation to calculate the flight load;

[0037] Step S204: Determine whether to fuel the space launch vehicle based at least on the flight payload envelope.

[0038] In traditional wind correction techniques, the mismatch between statistical and measured wind fields, coupled with conservative multi-stage trajectory designs, leads to excessive flight loads and low launch day release probabilities for spacecraft. These issues are mitigated in this embodiment. Figure 3 As shown, in the statistical wind field model, the diamond-shaped dashed line represents the statistical maximum value, the square dashed line represents the 75% statistical maximum value, the triangular dashed line represents the 50% statistical maximum value, the circular dashed line represents the 25% statistical maximum value, and the solid line represents the measured wind speed value. It can be seen that the wind direction and speed of the measured upper-level wind field differ significantly from those of the statistical wind field model. The measured upper-level wind field can accurately reflect the complex wind field characteristics such as transient wind shear and turbulence in actual flight, significantly improving the characterization accuracy of wind shear gradient and transient turbulence characteristics, and overcoming 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, making the load envelope closer to actual flight, avoiding the "margin superposition" of traditional methods, reducing aerodynamic angle of attack and flight load, increasing launch release probability, and expanding the adaptability of the launch window.

[0039] Unlike traditional four-level wind correction technology, quasi-real-time wind correction technology does not use statistical wind profiles in trajectory design. Instead, it directly uses pre-launch measured wind as the trajectory design input. Quasi-real-time wind correction technology can replace the previous statistical wind field, significantly improving the accuracy of wind field characterization and reducing angle-of-attack calculation errors from the source. When a spacecraft flies in a high-wind area with a trajectory corrected by quasi-real-time wind, the angle of attack is theoretically close to 0°, fundamentally reducing the spacecraft's flight load. Under ideal conditions, when the input high-altitude wind field data is the wind field when the spacecraft actually flies through a high-wind area, the aerodynamic angle of attack is 0°.

[0040] In practice, it is not possible to obtain the wind speed vector during the flight of a spacecraft in real time and design a corresponding wind-corrected trajectory. In this embodiment, measured upper-level wind data from a period of time before the launch window is selected. For example, measured upper-level wind data from one hour prior to the launch window is selected as the wind field for the launch window. Based on this, the upper-level wind-corrected trajectory design is carried out, i.e., quasi-real-time wind-corrected trajectory design. It can be seen that the closer the time of obtaining the upper-level wind data is to the launch window, the better the wind correction effect. Therefore, in the implementation process, reasonably selecting the time to measure the upper-level wind field is an important step in ensuring the quasi-real-time wind correction effect.

[0041] Alternatively, this embodiment can employ machine learning-assisted trajectory design, using historical wind field trajectory data to train a neural network, thereby achieving rapid optimization of trajectory correction parameters and further shortening the design time.

[0042] Specifically, the launch vehicle travels to the meteorological station to obtain upper-level wind data for the next three hours (-3 hours). This upper-level wind data is then transmitted back as input for trajectory correction. Based on this data, near-real-time wind correction design, simulation, analysis, and judgment are performed. Afterward, flight payloads are calculated using a six-degree-of-freedom Monte Carlo simulation, and the payload envelopes of each component of the launch vehicle are obtained. The propellant loading for the launch vehicle is determined based on at least the value of the payload 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.

[0043] Next, the launch vehicle travels to the meteorological station to obtain negative one-hour upper-level wind data. This data is then transmitted back as input for trajectory correction. Based on this negative one-hour upper-level wind data, near-real-time wind correction design, simulation, analysis, and judgment are performed. After six-degree-of-freedom Monte Carlo simulation, the flight payload is calculated, and the payload envelopes of each component of the launch vehicle are obtained. The launch vehicle's release is determined based at least on the values ​​of these payload envelopes. If the release criteria are met, near-real-time wind-corrected trajectory and guidance data are output. Upon entering the launch process, the negative one-hour launch data is transmitted to the launch vehicle control center. After checks, format conversions, and uploading, the guidance data is finally compiled. If the launch conditions are not met, the launch process is aborted.

[0044] 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 time requirements for decision-making on the launch day and supporting near real-time trajectory adjustment.

[0045] According to one embodiment of the present invention, the high-altitude wind data collected from the negative 1 hour and the negative 3 hours includes at least the following: the distribution data of wind speed and wind direction with altitude in the launch area.

[0046] According to one embodiment of the present invention, the method of using negative 1 hour upper-level wind data as input for ballistic correction and negative 3 hour upper-level wind data as input for ballistic correction specifically involves: solving the ballistic motion equation based on measured upper-level wind data with the minimum aerodynamic angle of attack as the objective, wherein the longitudinal axis of the space launch vehicle is aligned with the airspeed vector in the transonic range and the maximum dynamic pressure region, while simultaneously satisfying the orbital insertion condition constraints.

[0047] According to one embodiment of the present invention, the output guidance parameters are specifically generated as follows: based on the quasi-real-time wind-corrected trajectory, at least the program angle, perturbation guidance coefficient, standard guidance quantity, iterative guidance parameters, and shutdown equation coefficients are generated as guidance parameters.

[0048] According to one embodiment of the present invention, the method for six-degree-of-freedom Monte Carlo target simulation is as follows: injecting random errors, including at least inertial group noise and wind field deviation, in a normal distribution manner, and correcting them using simulated satellite navigation combined with navigation to simulate real flight conditions, thereby obtaining flight area safety, landing area safety, propellant reserve, and orbital insertion accuracy.

[0049] According to one embodiment of the present invention, the method for calculating flight loads specifically involves: extracting data on aerodynamic angle of attack and engine sway angle during the flight phase from a six-degree-of-freedom Monte Carlo shooting simulation method; calculating lateral normal aerodynamic loads, axial static loads, sway loads, gust loads, and flutter loads; and comprehensively obtaining the equivalent axial compression and axial tension load envelopes for each section of the rocket body.

[0050] According to one embodiment of the present invention, in addition to the flight payload envelope, the flight area safety, landing area safety, propellant reserve and orbital insertion accuracy must all meet the launch requirements.

[0051] According to one embodiment of the present invention, the flight area safety is that the flight trajectory in the six-degree-of-freedom Monte Carlo simulation does not exceed the safety conduit boundary; the landing area safety is that the first stage and fairing debris in the six-degree-of-freedom Monte Carlo simulation are located within the landing area.

[0052] According to one embodiment of the present invention, the propellant reserve is such that the probability of first-stage exhaustion is no greater than 15%, and the probability of second-stage exhaustion before orbit insertion is 0; the orbit insertion accuracy is such that the deviation of the semi-major axis of the orbit is no greater than 5 km, the deviation of the orbital eccentricity is no greater than 0.002, and the deviation of the orbital inclination angle is no greater than 0.08°.

[0053] Specifically, based on measured wind speed and direction distribution data with altitude, the ballistic equations are solved with the goal of minimizing the aerodynamic angle of attack, 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 range and the maximum dynamic pressure region, i.e., the theoretical angle of attack approaches 0°, while simultaneously satisfying orbital insertion constraints. After obtaining the quasi-real-time wind-corrected trajectory scheme based on the pre-launch measured wind field, guidance parameters are calculated, including the program angle, perturbation guidance coefficients, standard guidance quantity, iterative guidance parameters, and characteristic quantities of shutdown equation coefficients. All of these guidance parameters are designed automatically.

[0054] Next, to verify whether the near-real-time wind-corrected trajectory scheme and the guidance parameters and control parameters designed based on the near-real-time wind-corrected trajectory meet mission requirements, a six-degree-of-freedom Monte Carlo fire simulation verification was conducted. The fire simulation test randomly injected various errors according to a normal distribution, and the test introduced inertial navigation noise and an inertial navigation model, using simulated satellite navigation to simulate integrated navigation corrections. For example, in one embodiment, more than 2000 six-degree-of-freedom simulations can be conducted to simulate real flight conditions. Considering the impact of the impact area wind field on the landing point of the first stage and the fairing, the statistical high-altitude wind field of the impact area was used during the design phase, and the predicted wind field of the impact area one day prior to the launch date was used during mission execution. Through simulation verification, the guidance system correctly outputs commands, the attitude control system manipulates the servo mechanism, swings the engines, and controls the spacecraft to fly stably along the near-real-time wind-corrected trajectory.

[0055] The flight load calculation primarily targets the phase of the spacecraft traversing high-wind areas, specifically 30-110 seconds after liftoff. Data such as aerodynamic angle of attack and engine yaw angle are obtained from a six-DOF Monte Carlo simulation. Lateral normal aerodynamic loads, axial static loads, sloshing loads, gust loads, and flutter loads are calculated, comprehensively yielding the equivalent axial compression and tension load envelopes for each section of the launch vehicle. Axial static loads are determined by engine thrust, aerodynamic drag, and longitudinal overload. Lateral normal aerodynamic loads must consider the sloshing load of the liquid within the propellant tanks, the elastic response load caused by gust interference (i.e., gust load), and the pulsating pressure load in the transonic range (i.e., flutter load). By superimposing the static and dynamic loads at the same moment, the axial force, shear force, and bending moment of each section of the spacecraft are obtained, and the equivalent axial compression and tension are calculated. Load evaluation, based on Monte Carlo simulation results, quantifies load distribution characteristics and avoids conservative design.

[0056] Alternatively, this embodiment may employ multi-objective joint optimization technology to couple load reduction with indicators such as orbital accuracy and propellant consumption, and achieve comprehensive performance improvement through optimization tools such as genetic algorithms.

[0057] In practical operation, to apply near-real-time wind correction technology to determine whether the upper-altitude wind field before launch meets the launch requirements, corresponding release criteria need to be formulated. Release criteria consist of parameters and judgments, mainly considering flight area safety, landing area safety, propellant remaining quantity, orbital insertion accuracy, and flight payload factors. Specifically, flight area safety requires that the six-DOF Monte Carlo simulation flight trajectory not exceed the safety conduit boundary; landing area safety requires that the first stage and fairing debris be located within the landing area; propellant safety margin requires that the first stage exhaustion probability not exceed 15%, and the second stage exhaustion probability before orbital insertion be 0; satellite orbital insertion accuracy requires that the orbital semi-major axis deviation not exceed 5 km, the orbital eccentricity deviation not exceed 0.002, and the orbital inclination deviation not exceed 0.08°.

[0058] It should be noted that, prior to the launch vehicle's flight mission, based on the ultimate load-bearing capacity, safety factor, residual strength factor, and launch reliability requirements of each section of the launch vehicle's structural components, the permissible load-bearing limits for each structural section during flight in the first-level high-wind zone were determined. Based on pre-launch wind measurements, quasi-real-time wind-induced design and simulation work was conducted to predict the equivalent axial compression and equivalent axial tension load envelopes for each structural section during flight in the first-level high-wind zone. When the pre-launch predicted load envelope is less than the permissible load-bearing limit, the load release conditions are met.

[0059] The method described in this embodiment significantly improves the launch probability of space launch vehicles, increasing it from 95% using traditional techniques to 99.95%, thus reducing launch delay costs. This method integrates multidisciplinary models of wind field, trajectory, payload, and control, achieving closed-loop optimization from wind field measurement to launch decision-making, thereby improving system design efficiency. This method supports launches under more demanding wind field conditions, enhancing the mission flexibility of space launch vehicles and adapting to the high-frequency launch requirements of commercial spaceflight. In this embodiment, the reduced payload allows for a lighter design weight of the rocket body, thereby improving the overall payload capacity.

[0060] The above are merely preferred embodiments of the present invention and are not intended to limit the present invention. Any modifications, equivalent substitutions, improvements, etc., made within the spirit and principles of the present invention should be included within the protection scope of the present invention.

Claims

1. A method of pre-launch wind correction for a space launch vehicle, the method comprising: The method comprises the following steps: Collecting negative 1-hour measured high-altitude wind data, the high-altitude wind data at least comprising: collected wind speed and wind direction distribution data with height in the launch area; Taking the negative 1-hour high-altitude wind data as input of trajectory correction, and outputting guidance parameters; solving the trajectory motion equation based on the measured high-altitude wind data with the minimum aerodynamic attack angle as the target, wherein the longitudinal axis of the space vehicle is consistent with the airspeed vector in the transonic section and the maximum dynamic pressure area, while satisfying the orbit entry condition constraint; Developing a six-degree-of-freedom Monte Carlo targeting simulation to calculate flight load envelope; in the process of developing the six-degree-of-freedom Monte Carlo targeting simulation: injecting random errors of at least inertial measurement unit noise and wind field deviation in a normal distribution manner, simulating combined navigation correction by using a simulation satellite, to simulate real flight conditions, and obtaining flight area safety, landing area safety, propellant reserve and orbit entry accuracy; extracting the data of aerodynamic attack angle and engine swing angle in the flight phase of the six-degree-of-freedom Monte Carlo targeting simulation method, calculating lateral and normal aerodynamic loads, axial static loads, sway loads, gust loads and buffeting loads, and comprehensively obtaining equivalent axial pressure and axial tension load envelopes of each section of the rocket body; The axial static load is determined by the engine thrust, aerodynamic resistance and longitudinal overload, and the lateral and normal aerodynamic loads need to consider the sway load of the liquid in the tank, the elastic response load caused by gust interference, i.e. gust load, and the pulsating pressure load in the transonic section, i.e. buffeting load; superimposing the static and dynamic loads at the same time to obtain the axial force, shear force and bending moment of each section of the space vehicle, and then calculating the equivalent axial pressure and equivalent axial tension, and then calculating the equivalent axial pressure and axial tension load envelope of each section, which is the flight load envelope; Deciding whether to launch the space vehicle according to at least the flight load envelope.

2. The method of pre-launch wind correction for a space vehicle of claim 1, wherein, Before the step of collecting negative 1-hour measured high-altitude wind data, the method further comprises the following steps: Collecting negative 3-hour measured high-altitude wind data; Taking the negative 3-hour high-altitude wind data as input of trajectory correction, and outputting guidance parameters; Developing a six-degree-of-freedom Monte Carlo targeting simulation to calculate flight load; Deciding whether to fill the space vehicle according to at least the flight load envelope.

3. The method of pre-launch wind correction for a space vehicle of claim 2, wherein, The high-altitude wind data in the step of collecting negative 3-hour measured high-altitude wind data at least comprises: collected wind speed and wind direction distribution data with height in the launch area.

4. The method of pre-launch wind correction for a space vehicle of claim 3, wherein, The method of taking the negative 1-hour high-altitude wind data as input of trajectory correction and taking the negative 3-hour high-altitude wind data as input of trajectory correction specifically comprises: Solving the trajectory motion equation based on the measured high-altitude wind data with the minimum aerodynamic attack angle as the target, wherein the longitudinal axis of the space vehicle is consistent with the airspeed vector in the transonic section and the maximum dynamic pressure area, while satisfying the orbit entry condition.

5. The method of pre-launch wind correction for a space vehicle of claim 4, wherein, The output guidance parameters specifically comprise: Generating guidance parameters related to program angle, perturbation guidance coefficient, standard guidance amount, iterative guidance parameters and shutdown equation coefficient according to the quasi-real-time wind trajectory correction.

6. The method of pre-launch wind correction for a space vehicle of claim 5, wherein, The method of developing a six-degree-of-freedom Monte Carlo targeting simulation specifically comprises: Injecting random errors of at least inertial measurement unit noise and wind field deviation in a normal distribution manner, simulating combined navigation correction by using a simulation satellite, to simulate real flight conditions, and obtaining flight area safety, landing area safety, propellant reserve and orbit entry accuracy.

7. The method of pre-launch wind correction for a space vehicle of claim 6, wherein, The method for calculating flight load envelope, in particular: The method for extracting the data of the aerodynamic attack angle and the engine swing angle in the flight phase of the six-degree-of-freedom Monte Carlo targeting simulation, calculating the lateral normal aerodynamic load, the axial static load, the sway load, the gust load and the buffeting load, and comprehensively obtaining the equivalent axial pressure and the axial tension load envelope of each section of the missile body.

8. The method of pre-launch wind correction for a space vehicle of claim 6, wherein, In addition to the flight load envelope, the flight area safety, the landing area safety, the propellant surplus and the orbit accuracy all need to meet the launch requirements.

9. The method of pre-launch wind correction for a space vehicle of claim 8, wherein, The flight area safety is that the six-degree-of-freedom Monte Carlo targeting simulation flight trajectory does not exceed the safety pipeline boundary; the landing area safety is that the sub-stage and the fairing debris are located in the landing area in the six-degree-of-freedom Monte Carlo targeting simulation.

10. The method of pre-launch wind correction for a space vehicle of claim 8, wherein, The propellant surplus is that the sub-stage consumption probability is not greater than 15%, and the second stage consumption probability before entering the orbit is 0; the orbit accuracy is that the orbit semi-major axis deviation is not greater than 5 km, the orbit eccentricity deviation is not greater than 0.002, and the orbit inclination deviation is not greater than 0.08°.

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