A method for precise control of fairing landing point based on range separation

Abstract

This invention discloses a precise control method for fairing landing point based on range separation. It establishes dynamic models for the active and passive phases of rocket flight and a second-stage yaw program angle design model. Based on the active phase dynamic model, a range-based fairing jettison model is obtained. The fairing landing point is calculated based on the passive phase dynamic model, the second-stage yaw program angle design model, and the range-based fairing jettison model. The fairing landing point is corrected according to the environmental conditions at the landing point, resulting in a corrected fairing landing point. Through improved control methods, the landing area is effectively reduced; through improved trajectory optimization methods, the prominent contradiction between landing area requirements and orbital requirements is resolved; through improved target-firing methods, refined analysis and design of landing area data are achieved; and through historical sample analysis, examples and support for improved landing area design schemes are provided.

Description

Technical Field

[0001] This invention relates to a method for precise control of fairing landing point based on range separation, belonging to the field of launch vehicle trajectory design technology. Background Technology

[0002] With launch sites for carrier rockets located at inland bases, the issue of debris landing area control has become increasingly prominent. In recent years, with the development of the national economy and the rapid progress of production and construction in various regions, the selection of landing areas and the design of landing area ranges for launch missions at various inland launch bases have become increasingly difficult. In particular, for missions with certain specific launch azimuths, the landing areas of the rocket's first stage and fairing are distributed in densely populated areas in the southeast and along the coast, significantly compressing the range of landing area selection. Especially for fairing landing areas, where the flight trajectory of the passive phase is more uncertain, the requirements for refined design are becoming increasingly stringent, necessitating further research into landing area reduction design.

[0003] The need for controlling the fairing landing area is becoming increasingly urgent. On the one hand, economic development has led to a reduction in the available landing area within the country, making the selection and determination of new landing areas increasingly difficult. Surveying landing areas also requires coordinating human and material resources from the military, local governments, and scientific and technological departments, resulting in significant economic costs. On the other hand, after years of development, the pace of production and construction within existing mature landing areas is accelerating, leading to situations where mature landing areas become unusable due to the addition of high-value targets such as highways, railways, reservoirs, and dams. Therefore, it is imperative to optimize landing area design methods to reduce the landing area, thereby avoiding major adjustments to flight missions due to the lack of available landing areas, or reducing the carrying capacity of flight missions due to changes in landing area location. This will provide support for improving launch efficiency and reducing mission costs. Summary of the Invention

[0004] The technical problem solved by this invention is to overcome the shortcomings of the prior art and provide a precise control method for fairing landing point based on range separation. By improving the control method, the landing area is effectively reduced; by improving the trajectory optimization method, the prominent contradiction between landing area requirements and trajectory requirements is resolved; by improving the target shooting method, the landing area data can be refined for analysis and design; and by analyzing historical samples, the landing area design improvement scheme can be illustrated and supported.

[0005] The technical solution of this invention is:

[0006] This invention discloses a method for precise control of fairing landing point based on range separation, comprising the following steps:

[0007] Establish a dynamic model of the active segment;

[0008] Establish a dynamic model for the passive segment;

[0009] Establish a two-stage yaw procedure angle design model;

[0010] Based on the active segment dynamics model, the range-deployment model is obtained;

[0011] The fairing range is calculated based on the passive section dynamics model, the second-stage yaw procedure angle design model, and the range fairing jetting model.

[0012] Based on the fairing range and the environmental conditions at the impact point, the fairing impact point is corrected to obtain the corrected fairing range.

[0013] Furthermore, in the above control method, the active segment dynamic model is specifically as follows:

[0014]

[0015]

[0016] in, For the resultant acceleration of the inertial frame, x a y a z a For the position vector in the inertial frame, g ax g ay g az To generate gravitational acceleration in the inertial frame, To generate the apparent acceleration under the inertial frame, V ax V ay V az The velocity is the speed under the inertial frame.

[0017] Furthermore, in the above control method, the apparent acceleration of the inertial system is given by the following formula:

[0018]

[0019]

[0020] in, To generate the downward-looking acceleration of the inertial frame; For the downward-looking acceleration of the arrow system; F ax1 F ay1 F az1 Aerodynamics of the arrow system; P x1 P y1 P z1 For the engine thrust in the arrow system; ψ a Let m be the attitude angle under the inertial frame, and m be the mass of the rocket.

[0021] Furthermore, in the above control method, the gravitational acceleration under the inertial frame is defined by the following formula:

[0022]

[0023]

[0024]

[0025] Among them, g ax g ay g az To generate gravitational acceleration in the inertial frame, x a y a z a Here, A0 is the position vector in the inertial frame, and B0 is the geodetic azimuth and geodetic latitude. Where is the geocentric latitude, M is the Earth's mass, J2 is the Earth's zonal harmonic coefficient, G is the Earth's gravitational constant, and a e R is the average radius of the Earth's equator. ox R oy R oz Let r be the geocentric radius vector at a point on the ground, and g be the modulus of the geocentric radius vector. r g w This is the projection of gravitational acceleration onto the Earth's radius vector and tangential direction.

[0026] Furthermore, in the above control method, the passive segment dynamic model is specifically as follows:

[0027]

[0028]

[0029]

[0030] q=ρV 2 / 2

[0031]

[0032]

[0033] in, For the inertial frame to generate the combined acceleration, V ax V ay V az For the velocity in the inertial frame, w x w y w z To generate the local wind field vector in the inertial frame, C xlr g is the equivalent aerodynamic drag coefficient. ax g ay g az To generate gravitational acceleration in the inertial frame, k rHere, S is a dimensionless coefficient, S is the rocket's reference area, m is the rocket's mass, and C is a dimensionless coefficient. xlr Here, q is the equivalent aerodynamic drag coefficient, q is the flight kinematic pressure, g is the acceleration due to gravity, and n is the gravitational acceleration. x For axial overload, k r ρ is a dimensionless coefficient, and ρ0 is the corresponding altitude and the atmospheric density at 0 altitude.

[0034] Furthermore, in the above control method, the velocity under the inertial frame is defined by the following formula:

[0035] V ax =V ax +(y a +R 0y )ω z -(z a +R 0z )ω y

[0036] V ay =V ay +(z a +R 0z )ω x -(x a +R 0x )ω z

[0037] V az =V az +(x a +R 0x )ω y -(y a +R 0y )ω x

[0038] Among them, V ax V ay V az For the velocity in the inertial frame, ω x ω y ω z For the component of Earth's rotation in the inertial frame, x a y a z a R is the position vector in the inertial frame. ox R oy R oz Let be the geocentric radius of a point on the ground.

[0039] Furthermore, in the above control method, the second-level yaw procedure angle design model is specifically as follows:

[0040]

[0041] Where, ψpr The yaw program angle under the inertial frame of reference is given, t is the second-stage solo flight time, t0 to t5 are the six flight time points within time t, and ψ is... pr0 ψ is the initial yaw procedure angle for the second stage of flight. pr1 ψ is the yaw program angle during fairing jettison. pr2 This is the final yaw procedure angle for the second stage of flight; For ψ pr1 The corresponding rate of change of angle; For ψ pr2 The corresponding rate of change of angle.

[0042] Furthermore, in the above control method, the step of obtaining the range-deploying model based on the active segment dynamics model specifically involves:

[0043]

[0044] Where L represents the predicted rocket range, and V ax V ay V az For the velocity in the inertial frame, x a y a z a This is the position vector in the inertial frame.

[0045] Furthermore, in the above control method, the calculation of the fairing range based on the passive phase dynamics model, the second-stage yaw procedure angle design model, and the range fairing jettison model specifically involves:

[0046] L l =R m cos -1 {[(x l +R 0x )·R 0x +(y l +R 0y )·R 0y +(z l +R 0z )·R 0z ] / (r l ·R0)}

[0047] Among them, L l R represents the actual range of the rocket. m r is the average radius of the Earth. l x is the radius vector of the landing point. l y l z l Let R0 and R be the radial components of the impact point in the launching system. 0x R 0y R 0z This represents the geocentric radius vector corresponding to the launch point and its projection onto the launch system.

[0048] Furthermore, in the above control method, the local wind field vector under the inertial frame is obtained according to "GJB365.1-87 Standard Atmosphere for the Northern Hemisphere (-2 to 80 km)".

[0049] The advantages of this invention over the prior art are as follows:

[0050] (1) This invention achieves a significant reduction in lateral velocity and position dispersion by adding full perturbation guidance functionality to the sub-stage flight. The mechanism is to effectively track and limit lateral deviations by guiding the lateral range of a sub-stage laterally.

[0051] (2) Based on this range separation fairing landing point precision control method, the present invention can effectively reduce the sub-stage area by more than 50% for missions with tight sub-stage debris landing areas. Attached Figure Description

[0052] Figure 1 This is a target hit statistical chart showing the velocity and position deviation of the first and second separation points according to the present invention;

[0053] Figure 2 This is a comparison diagram of the lateral landing point distribution of the first-level guided and unguided objects of the present invention.

[0054] Figure 3 This is a block diagram illustrating the basic principle of the perturbation guidance of this invention;

[0055] Figure 4 This is a diagram of the fairing landing area of ​​the present invention;

[0056] Figure 5 This is a scatter diagram of the fairing landing points before and after the fairing throw control based on the present invention. Detailed Implementation

[0057] The present invention will be further described in detail below with reference to the accompanying drawings and specific embodiments.

[0058] First, the dynamics and guidance control equations involved in the analysis process are presented. The effects of first-stage guidance, range fairing jetting, and local wind field on the fairing landing area and landing point deviation are analyzed, and relevant technical improvements are made. Finally, the conclusions are supported by comprehensive deviation target shooting, thereby supporting the subsequent landing area control work of the model.

[0059] The impact area reduction is considered using both lateral and longitudinal paths. Firstly, based on the perturbation guidance principle and programmed angle tracking, a guidance design function is added to significantly reduce the lateral deviation of the rocket fairing debris impact point, as shown in the attached diagram. Figure 1 and attached Figure 2As shown, to eliminate excessive longitudinal range deviation of stage debris caused by exhaustion shutdown, the traditional time-course fairing design was changed to a range-course fairing design. This fully utilizes secondary capabilities and guidance to reduce longitudinal and lateral deviations, and design improvements were made using segmented piecewise linear design of the yaw program angle. Furthermore, the local wind field near the impact area was used instead of the traditional mean westerly wind as the passive section design wind field. Finally, multiple target tests were conducted on the fairing impact area to further guide the refined design of the fairing impact area and ensure the reliability of the design results.

[0060] This invention provides a method for precise control of fairing landing point based on range separation, comprising the following steps:

[0061] Step 1: Establish the dynamic model of the active segment;

[0062] Step 2: Establish the dynamic model of the passive segment;

[0063] Step 3: Establish the second-level yaw procedure angle design model;

[0064] Step 4: Based on the dynamic model of the active phase, obtain the range-deployment model;

[0065] Step 5: Calculate the fairing range based on the passive section dynamics model, the second-stage yaw procedure angle design model, and the range fairing jettison model;

[0066] Step 6: Based on the fairing range and the environmental conditions at the impact point, correct the fairing impact point to obtain the corrected fairing range.

[0067] Preferably, the active segment dynamics model is as follows:

[0068]

[0069]

[0070] in, For the resultant acceleration of the inertial frame, x a y a z a For the position vector in the inertial frame, g ax g ay g az To generate gravitational acceleration in the inertial frame, To generate the apparent acceleration under the inertial frame, V ax V ay V az The velocity is the speed under the inertial frame.

[0071] Preferably, the apparent acceleration under the inertial frame is given by the following formula:

[0072]

[0073]

[0074] in, To generate the downward-looking acceleration of the inertial frame; For the downward-looking acceleration of the arrow system; F ax1 F ay1 F az1 Aerodynamics of the arrow system; P x1 P y1 P z1 For the engine thrust in the arrow system; ψ a Let m be the attitude angle under the inertial frame, and m be the mass of the rocket.

[0075] Preferably, the gravitational acceleration in the inertial frame is given by the following formula:

[0076]

[0077]

[0078]

[0079] Among them, g ax g ay g az To generate gravitational acceleration in the inertial frame, x a y a z a Here, A0 is the position vector in the inertial frame, and B0 is the geodetic azimuth and geodetic latitude. Where is the geocentric latitude, M is the Earth's mass, J2 is the Earth's zonal harmonic coefficient, G is the Earth's gravitational constant, and a e R is the average radius of the Earth's equator. ox R oy R oz Let r be the geocentric radius vector at a point on the ground, and g be the modulus of the geocentric radius vector. r g w This is the projection of gravitational acceleration onto the Earth's radius vector and tangential direction.

[0080] Preferably, the passive segment dynamics model is as follows:

[0081]

[0082]

[0083]

[0084] q=ρV 2 / 2

[0085]

[0086]

[0087] in, For the inertial frame to generate the combined acceleration, V ax V ay V az For the velocity in the inertial frame, w x w y w z To generate the local wind field vector in the inertial frame, C xlr g is the equivalent aerodynamic drag coefficient. ax g ay g az To generate gravitational acceleration in the inertial frame, k r Here, S is a dimensionless coefficient, S is the rocket's reference area, m is the rocket's mass, and C is a dimensionless coefficient. xlr Here, q is the equivalent aerodynamic drag coefficient, q is the flight kinematic pressure, g is the acceleration due to gravity, and n is the gravitational acceleration. x For axial overload, k r ρ is a dimensionless coefficient, and ρ0 is the corresponding altitude and the atmospheric density at 0 altitude.

[0088] Preferably, the velocity in the inertial frame is given by the following formula:

[0089] V ax =V ax +(y a +R 0y )ω z -(z a +R 0z )ω y

[0090] V ay =V ay +(z a +R 0z )ω x -(x a +R 0x )ω z

[0091] V az =V az +(x a +R 0x )ω y -(y a +R 0y )ω x

[0092] Among them, V ax V ay V az For the velocity in the inertial frame, ωx ω y ω z For the component of Earth's rotation in the inertial frame, x a y a z a R is the position vector in the inertial frame. ox R oy R oz Let be the geocentric radius of a point on the ground.

[0093] Preferably, the second-stage yaw procedure angle design model is as follows:

[0094]

[0095] Where, ψ pr The yaw program angle under the inertial frame of reference is given, t is the second-stage solo flight time, t0 to t5 are the six flight time points within time t, and ψ is... pr0 ψ is the initial yaw procedure angle for the second stage of flight. pr1 ψ is the yaw program angle during fairing jettison. pr2 This is the final yaw procedure angle for the second stage of flight; For ψ pr1 The corresponding rate of change of angle; For ψ pr2 The corresponding rate of change of angle.

[0096] Preferably, based on the active phase dynamics model, the range-based missile launch model is obtained, specifically as follows:

[0097]

[0098] Where L represents the predicted rocket range, and V ax V ay V az For the velocity in the inertial frame, x a y a z a This is the position vector in the inertial frame.

[0099] Preferably, the fairing range is calculated based on the passive phase dynamics model, the second-stage yaw procedure angle design model, and the range jetting model, specifically as follows:

[0100] L l =R m cos -1 {[(x l +R 0x )·R 0x +(y l +R 0y )·R 0y +(z l +R 0z )·R0z ] / (r l ·R0)}

[0101] Among them, L l R represents the actual range of the rocket. m r is the average radius of the Earth. l x is the radius vector of the landing point. l y l z l Let R0 and R be the radial components of the impact point in the launching system. 0x R 0y R 0z This represents the geocentric radius vector corresponding to the launch point and its projection onto the launch system.

[0102] Preferably, the local wind field vector under the inertial frame is obtained according to "GJB365.1-87 Standard Atmosphere for the Northern Hemisphere (-2 to 80 km)".

[0103] Example

[0104] This embodiment provides a method for precise control of fairing landing point based on range separation. The specific implementation method and steps are as follows:

[0105] The first step is to determine the dynamic equations of the rocket's active phase.

[0106] The rocket's active phase trajectory calculation model is as follows

[0107]

[0108]

[0109] The formula for calculating apparent acceleration is:

[0110]

[0111]

[0112] To generate the downward-looking acceleration of the inertial frame;

[0113] For the downward-looking acceleration of the arrow system;

[0114] F ax1 F ay1 F az1 Aerodynamics of the arrow system;

[0115] P x1 P y1 P z1 The thrust of the engine in the arrow system.

[0116] The formula for calculating gravitational acceleration is:

[0117]

[0118]

[0119]

[0120] The second step is to determine the dynamic equations of the rocket's passive phase.

[0121] During reentry, the substage and fairing are no longer under control, meaning there are no control force-related terms. Without control forces, the substage exhibits attitude roll characteristics at different times during reentry. In engineering, considering it as a point mass only subject to aerodynamic forces and weight, the passive phase dynamic equations are as follows.

[0122]

[0123]

[0124] V ax =V ax +(y a +R 0y )ω z -(z a +R 0z )ω y

[0125] V ay =V ay +(z a +R 0z )ω x -(x a +R 0x )ω z

[0126] V az =V az +(x a +R 0x )ω y -(y a +R 0y )ω x

[0127]

[0128] q=ρV 2 / 2

[0129]

[0130]

[0131] The third step is to improve the ballistic design method.

[0132] The improvement in design methodology mainly involves the design of the second-stage yaw program angle. Previously, rockets typically employed a single-segment linear polygonal design for the yaw path during the second-stage flight phase.

[0133]

[0134] The above methods typically achieve a monotonically linear change in the yaw program angle during the first half of the second-stage flight phase, and maintain the yaw program angle during the second half. Designing the yaw angle during the second-stage flight phase results in a loss of payload capacity, and its initial design purpose is to accommodate landing area constraints. However, for certain special missions, the traditional monotonically linear design mode for the constant-value flight phase yaw program angle is no longer sufficient. Yaw design in the opposite orbital inclination direction is also required to meet the fairing landing area constraints on fairing jettison conditions. Therefore, the rocket will fly in an arc during the second-stage flight phase, thus satisfying both the satellite's orbital inclination requirements and the fairing landing area constraints.

[0135] Fourth step, control method improvement

[0136] The fairing landing area formed by the traditional timing-based fairing ejection, which is based on scheduled shutdown, has a large trailing edge dispersion. This is because it is greatly affected by the first-stage exhaustion shutdown.

[0137] Therefore, it is proposed to modify the timing-based fairing ejection method to a range-based method, thereby effectively reducing the longitudinal range dispersion of the fairing landing area. The mathematical model for range-based fairing ejection is as follows.

[0138]

[0139] The purpose of employing a range-based fairing ejection method is to control the longitudinal dispersion range of the fairing debris impact point. Because the fairing experiences significant air resistance during reentry and free flight, the impact of aerodynamic drag on the debris impact range must be considered when calculating the range partial derivative.

[0140] In calculating the relative velocity (V) of the range x V y V z When calculating the partial derivatives of position (X, Y, Z) and time (t), the partial derivatives of the relative velocity, position, and time of the first-stage rocket debris impact point are calculated starting from the velocity, position, and time at the theoretical shutdown moment. This yields the range ejection quantity. Lateral range dispersion control is ensured by incorporating a guidance function. The relevant mathematical model for first-stage guidance is attached. Figure 3 As shown.

[0141] When calculating the impact range, the ballistic calculation equation terminates by controlling the integration at the impact elevation. At this point, the radius vector from the Earth's center to the impact point can be obtained. This impact radius vector is represented in the launch coordinate system. Therefore, the impact radius vector can be transformed into a specific position in the launch coordinate system, and the impact range can be further calculated.

[0142] Fifth step: Improvement of target shooting methods

[0143] The depletion of propellant during the first stage shutdown has a significant impact on the safety margin of the second stage. Especially in certain launch missions with tight capacity, it is necessary to carry out refined design and analyze the safety margin of the second stage under the condition that the first stage is not depleted, so as to provide theoretical support for ensuring the first stage guidance shutdown. As a result, a lot of improvements were made to the design and simulation program and the following work was carried out: (1) the initial conditions of the first stage depletion shutdown were eliminated and the second stage depletion situation was statistically analyzed; (2) the wind field of the landing area was corrected according to the national standard. When performing the passive segment calculation in the atmosphere, after the fairing flight altitude is below 80km, the wind speed and wind direction vector model is added according to the national standard to correct the actual landing point position, and the impact of different wind fields on the target hit results during the reentry segment is compared based on this method.

[0144] Following the steps outlined above, perform 5000 Monte Carlo simulations. The distribution of fairing debris impact points is shown in the appendix. Figure 4 Based on 3σ probability statistics, the fairing's longitudinal range is approximately 70 km, and its lateral range is approximately 30 km. Compared to the existing landing area, the longitudinal range is reduced by 12.5%, and the lateral range is reduced by 60%.

[0145] Based on this, a range-based fairing ejection analysis was conducted under first-level guidance. This analysis considered the random and uncontrollable attitude reversal of the fairing during the passive phase, and rigorously accounted for aerodynamic coefficient deviations during the passive phase (random deviations during target engagement were amplified to 90%). Furthermore, 5000 Monte Carlo target engagement simulations were performed, yielding the impact point dispersion under both time-based and range-based fairing ejection conditions, as shown in the attached figure. Figure 5 As shown in the figure. It can be seen that even with amplified passive section deviation, range fairing jetting can still effectively reduce the impact of near-miss caused by exhaustion shutdown, significantly reducing the fairing's trailing phenomenon within the longitudinal range, thus contributing to a smaller impact area. (See attached figure.) Figure 5 It is evident that by adopting range jetting, the fairing's longitudinal distance can be further reduced by more than 50% from the current level.

[0146] At this point, the calculations, design, and simulations related to the precise control technology for fairing landing points based on range separation have all been completed.

[0147] Although the present invention has been described in detail through the preferred embodiments above, it should be understood that the above description should not be considered as a limitation of the present invention. Various modifications and substitutions to the present invention will be apparent to those skilled in the art after reading the above description. Therefore, the scope of protection of the present invention should be defined by the appended claims.

[0148] The contents not described in detail in this specification are common knowledge to those skilled in the art.

Claims

1. A method for precise control of fairing landing point based on range separation, characterized in that, Includes the following steps: Establish a dynamic model of the active segment; Establish a dynamic model for the passive segment; Establish a two-stage yaw procedure angle design model; Based on the active segment dynamics model, the range-deployment model is obtained; The fairing range is calculated based on the passive section dynamics model, the second-stage yaw procedure angle design model, and the range fairing jetting model. Based on the fairing range and the environmental conditions at the impact point, the fairing impact point is corrected to obtain the corrected fairing range; The second-level yaw procedure angle design model is as follows: in, To generate the yaw procedure angle under the inertial frame, For Level 2 solo flight time, for Six flight time points within the time period, This is the initial yaw procedure angle for the second stage of flight. The yaw program angle during fairing jettison. This is the final yaw procedure angle for the second stage of flight; for The corresponding rate of change of angle; for The corresponding rate of change of angle; The range-deploying model obtained based on the active segment dynamics model is as follows: in, To indicate the rocket's range, , , For the speed under the inertial frame, The position vector in the inertial frame; The fairing range is calculated based on the passive section dynamics model, the second-stage yaw procedure angle design model, and the range fairing jettison model, specifically as follows: in, This refers to the actual range of the rocket. The average radius of the Earth For the radius of the landing point, Let the radial component of the impact point be in the launching frame. This represents the geocentric radius vector corresponding to the launch point and its projection onto the launch system.

2. The method for precise control of fairing landing point based on range separation according to claim 1, characterized in that: The active segment dynamics model is as follows: in, To generate the combined acceleration of the inertial frame, The derivative of the position vector in the inertial frame. To generate gravitational acceleration in the inertial frame, To generate inertial frame downward acceleration, , , The velocity is the speed under the inertial frame.

3. The method for precise control of fairing landing point based on range separation according to claim 2, characterized in that: The formula for the apparent acceleration of the inertial system is: in, To generate the downward-looking acceleration of the inertial frame; For the downward-looking acceleration of the arrow system; Aerodynamics of the arrow system; For the engine thrust in the arrow system; To generate the attitude angle under the inertial frame, For rocket mass.

4. The method for precise control of fairing landing point based on range separation according to claim 2, characterized in that: The formula for the gravitational acceleration in the inertial frame is: in, To generate gravitational acceleration in the inertial frame, The position vector in the inertial frame, For the azimuth angle of the large area, Latitude of the earth The latitude of the Earth's core. For Earth mass, The Earth's harmonic coefficients, The gravitational constant of Earth, The average radius of the Earth's equator. Let the geocentric radius be the point on the ground. For the modulus of the geocentric radius, This is the projection of gravitational acceleration onto the Earth's radius vector and tangential direction.

5. The method for precise control of fairing landing point based on range separation according to claim 1, characterized in that: The passive segment dynamics model is as follows: in, To generate the combined acceleration of the inertial frame, , , For the speed under the inertial frame, To generate the local wind field vector under the inertial frame, To generate gravitational acceleration in the inertial frame, Here, S is a dimensionless coefficient, S is the rocket's reference area, and m is the rocket's mass. This is the equivalent aerodynamic drag coefficient. For flight motion pressure, It is the acceleration due to gravity. For axial overload, The coefficient is dimensionless. This represents the atmospheric density at the corresponding altitude and at 0 altitude. It is the derivative of the position vector in the inertial frame.

6. The method for precise control of fairing landing point based on range separation according to claim 5, characterized in that: The velocity in the inertial frame is calculated using the following formula: in, , , For the speed under the inertial frame, The component of Earth's rotation in the inertial frame. The position vector in the inertial frame, Let be the geocentric radius of a point on the ground.

7. The method for precise control of fairing landing point based on range separation according to claim 5, characterized in that: The local wind field vector under the inertial frame was obtained according to "GJB365.1-87 Standard Atmosphere for the Northern Hemisphere (-2 to 80 km)".

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