A launch vehicle high-altitude wind offload control method and device

Abstract

The application provides a high-altitude wind load reduction control method and device for a carrier rocket, and the method comprises the following steps: obtaining a lateral normal apparent velocity increment of a rocket body coordinate system, and obtaining a lateral normal overload according to the lateral normal apparent velocity increment; integrating the lateral normal overload to obtain a lateral normal velocity increment; performing amplitude limiting processing and linear transition processing on the lateral normal velocity increment to obtain a compensation instruction value; obtaining a final rudder instruction value according to the sum of a rudder instruction value obtained by an attitude control system of the carrier rocket and the compensation instruction value; and controlling the carrier rocket by an execution mechanism of the carrier rocket, and the carrier rocket moves in an attitude related to the final rudder instruction value. The application can reduce the aerodynamic load of the carrier rocket, thereby improving the high-altitude wind field load adaptability, reducing the structural weight of the rocket, improving the carrying capacity, and enhancing the task-free ability.

Description

Technical Field

[0001] This application relates to the field of aerodynamic load control technology for launch vehicles, specifically to a method and device for high-altitude wind load reduction control of launch vehicles. Background Technology

[0002] During the high-altitude flight phase (5km–20km), launch vehicles encounter complex weather conditions. Under the influence of wind fields, the rocket's angle of attack (sideslip angle) is large, resulting in significant aerodynamic loads. To accommodate these uncertain loads, rocket structural designers often need to determine structural strength based on the wind load generated by the angle of attack within the wind envelope (which can be called the wind angle of attack). This significantly increases the rocket's structural mass, making it overly heavy and limiting its payload capacity. High-altitude wind aerodynamic loads restrict further improvements in rocket performance and increase development costs.

[0003] The currently used general load reduction method is ballistic correction load reduction technology based on pre-launch statistical wind field data. This involves introducing pre-measured high-altitude stable wind (statistical wind) information into the control system and using an open-loop correction method to compensate for the wind load. This open-loop compensation scheme does not change the original control structure and is simple and reliable. However, this scheme has the following shortcomings:

[0004] 1. When the actual wind deviates significantly from the statistical wind, the aerodynamic angle of attack is larger and the linearity of aerodynamic parameters deteriorates. Even when the deviation of aerodynamic parameters is small, it is still impossible to eliminate the large aerodynamic load caused by the difference between the two. When the aerodynamic parameters are pulled to their limits, the attitude of the rocket body is also prone to divergence. Excessive loads place high demands on the structural strength of the rocket body and the quality of the rocket body materials.

[0005] 2. It has a weak ability to adapt to changes in the upper-level wind field. When shear wind occurs, the system response speed may not be able to keep up with the occurrence of short-term large angle of attack and large overload phenomena.

[0006] 3. When air and wind conditions are poor, timely launch is not possible, and the requirement for de-tasking of launch vehicles cannot be met. Summary of the Invention

[0007] The purpose of this application is to provide a method and device for high-altitude wind load reduction control of launch vehicles, which can reduce aerodynamic load, improve the load adaptability of high-altitude wind fields, and enhance the ability to be de-tasked.

[0008] To achieve the above objectives, firstly, this application provides a method for high-altitude wind load reduction control of a launch vehicle, comprising:

[0009] Obtain the lateral normal velocity increment in the rocket body coordinate system, and obtain the lateral normal overload based on the lateral normal velocity increment;

[0010] Integrating the transverse normal overload yields the transverse normal velocity increment;

[0011] The transverse normal velocity increment is subjected to amplitude limiting and linear transition processing to obtain the compensation command value;

[0012] The final rudder command value is obtained by summing the rudder command value and the compensation command value obtained from the attitude control system of the launch vehicle. The actuators of the launch vehicle control the launch vehicle to move in attitude associated with the final rudder command value.

[0013] In some embodiments of this application, based on the aforementioned scheme, the lateral normal velocity increment of the rocket body coordinate system is obtained, and the lateral normal overload is obtained based on the lateral normal velocity increment, including:

[0014] The lateral normal velocity increment is obtained by combining inertial measurement and control data, and the lateral normal overload is obtained according to the first formula; the lateral normal overload includes lateral overload and normal overload, and the lateral normal velocity increment includes lateral velocity increment and normal velocity increment.

[0015] The first equation is:

[0016]

[0017]

[0018] In the formula, Ny represents the normal overload, and ΔW y1 G0 is the normal apparent velocity increment, G0 is the equivalent gravitational acceleration, Nz is the lateral overload, and ΔW is the apparent velocity increment. z1 The lateral apparent velocity increment is given by t, which is the preset time.

[0019] In some embodiments of this application, based on the aforementioned scheme, t is 0.01s.

[0020] In some embodiments of this application, based on the foregoing scheme, before the step of integrating the lateral normal overload to obtain the lateral normal velocity increment, the method further includes:

[0021] Filtering is applied to the transverse normal overload.

[0022] In some embodiments of this application, based on the aforementioned scheme, the transfer function of the filter used for filtering transverse normal overload is:

[0023]

[0024] Where T0 is the time constant.

[0025] In some embodiments of this application, based on the aforementioned scheme, the integral function for integrating the transverse normal overload is:

[0026] C(i) = k·T·R(i) + C(i-1)

[0027] Where k is the integral gain, T is the launch vehicle control cycle, C(i) is the output of the current launch vehicle control cycle, C(i-1) is the output of the previous launch vehicle control cycle, and R(i) is the input of the current launch vehicle control cycle.

[0028] In some embodiments of this application, based on the aforementioned scheme, k ranges from 2 to 3, and T is 0.01s.

[0029] In some embodiments of this application, based on the aforementioned scheme, the transverse normal velocity increment is subjected to amplitude limiting and linear transition processing to obtain a compensation command value, including:

[0030] The launch vehicle's angle of attack and flight altitude are obtained through the attitude control system;

[0031] The compensation command value is obtained based on the transverse normal velocity increment and the second formula;

[0032] The second equation is:

[0033]

[0034]

[0035] δ2=k b ·δ0

[0036] In the formula, δ p α is the lateral normal velocity increment, h1 is the launch vehicle's angle of attack, h2 is the altitude, h is the preset altitude, H is the launch vehicle's flight altitude, and δ2 is the compensation command value.

[0037] In some embodiments of this application, based on the aforementioned scheme, h1 is 2500m and h2 is 20000m.

[0038] Secondly, this application provides a high-altitude wind load reduction control device for a launch vehicle, comprising:

[0039] The first calculation unit is used to obtain the horizontal normal velocity increment of the rocket body coordinate system and to obtain the horizontal normal overload based on the horizontal normal velocity increment.

[0040] The integrating unit is used to integrate the lateral normal overload to obtain the lateral normal velocity increment;

[0041] The processing unit is used to perform amplitude limiting and linear transition processing on the transverse normal velocity increment to obtain the compensation command value.

[0042] The second calculation unit is used to obtain the final rudder command value based on the sum of the rudder command value and the compensation command value obtained from the attitude control system of the launch vehicle; the actuator of the launch vehicle controls the launch vehicle to move in attitude associated with the final rudder command value.

[0043] This application provides a method and apparatus for high-altitude wind load reduction control of a launch vehicle. The method includes: acquiring the lateral normal apparent velocity increment in the launch vehicle coordinate system; obtaining the lateral normal overload based on the lateral normal apparent velocity increment; integrating the lateral normal overload to obtain the lateral normal velocity increment; performing amplitude limiting and linear transition processing on the lateral normal velocity increment to obtain a compensation command value; obtaining the final rudder command value based on the sum of the rudder command value and the compensation command value obtained from the launch vehicle's attitude control system; and controlling the launch vehicle's actuators with the attitude motion command value associated with the final rudder command value. This allows the launch vehicle to fly with the wind, reducing the angle of attack and achieving the goal of reducing aerodynamic loads, thereby solving the high-altitude wind load problem and further improving the high-altitude wind field load adaptability. Furthermore, it eliminates the need for pre-launch wind field configuration, enhancing the rocket's task-independent capability. Attached Figure Description

[0044] The accompanying drawings, which are incorporated in and form part of this specification, illustrate embodiments consistent with this application and, together with the description, serve to explain the principles of this application. It is obvious that the drawings described below are merely some embodiments of this application, and those skilled in the art can obtain other drawings based on these drawings without any inventive effort.

[0045] In the attached diagram:

[0046] Figure 1 This is a schematic diagram of the conventional arrow body attitude control principle;

[0047] Figure 2 This is a schematic diagram illustrating the control principle of small angle of attack and low load in an embodiment of this application.

[0048] Figure 3 This is a flowchart of the high-altitude wind load reduction control method for launch vehicles according to an embodiment of this application;

[0049] Figure 4 This is a schematic diagram of the structure of the high-altitude wind load reduction control device for a launch vehicle according to an embodiment of this application. Detailed Implementation

[0050] To make the objectives, technical solutions, and advantages of this application clearer, the technical solutions in the embodiments of this application will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only a part of the embodiments of this application, and not all of the embodiments. Based on the embodiments in this application, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of this application.

[0051] The following detailed description of some embodiments of this application will be provided in conjunction with the accompanying drawings. Unless otherwise specified, the following embodiments and features can be combined with each other.

[0052] During actual flight, the wind field changes drastically at altitudes of 5km to 20km, with maximum wind speeds reaching 73m / s and angles of attack α (sideslip angle) reaching approximately 3.5°. Furthermore, during rocket flight, the dynamic pressure Q reaches its maximum at altitudes of approximately 7km to 15km, resulting in the largest Q·α value and significant aerodynamic loads, which can cause the rocket body to bend. Therefore, aerodynamic load control is necessary during rocket flight.

[0053] like Figure 1 As shown, the conventional attitude control principle of the rocket body is as follows: a control scheme of "strap-inertial navigation system - rocket computer - actuator" is adopted. The inertial navigation system (a combination of gyroscope and accelerometer, referred to as inertial navigation system) fixed to the rocket body outputs the apparent velocity increment and angular increment in the coordinate axis direction. After navigation calculation, the quaternion of the rocket body attitude motion is obtained. It is compared with the programmed attitude angle to form an angular deviation signal. After the attitude control system calculates the control command value, it drives the servo motor to rotate the corresponding angle, generating control force and control torque, which are applied to the rocket body, thereby manipulating the attitude motion of the rocket body, keeping the rocket body attitude stable and realizing the tracking of the programmed attitude angle.

[0054] The rocket's control loop mainly consists of two parts: attitude control and low-angle-of-attack, low-load control. Attitude control includes actual attitude angle measurement, angle deviation calculation, filter calculation, correction network calculation, and output of rudder command value δ1, which is a conventional control method.

[0055] This application focuses on the small angle-of-attack and low-load control portion, such as... Figure 2As shown, the principle of low angle-of-attack and low load control is as follows: Ignoring engine skew interference and tangential acceleration, the Y1 overload sensed by the inertial navigation system (INS) can be approximated as the result of aerodynamic forces generated by the wind angle of attack; the Z1 overload sensed by the INS can be approximated as the result of aerodynamic forces generated by the wind sideslip angle. At this point, the overload is calculated based on the transverse normal acceleration output from the accelerator output. After filtering and integration, the transverse normal velocity increment is obtained. This increment is then compensated to the actuator (servo motor) output after nonlinear limiting adjustment, allowing the rocket body to fly with the wind, reducing the aerodynamic angle of attack (sideslip angle), and thus reducing the aerodynamic load.

[0056] Based on this, such as Figure 3 As shown in the figure, this application provides a method for controlling high-altitude wind load reduction of a launch vehicle, including the following steps:

[0057] Step S1: Calculation of lateral normal overload. Obtain the lateral normal apparent velocity increment in the rocket body coordinate system, and calculate the lateral normal overload based on the lateral normal apparent velocity increment;

[0058] Specifically, the inertial measurement unit outputs information from the gyroscope and accelerometer every preset time t, for example, 10ms, and outputs the lateral normal apparent velocity increment of the arrow system accordingly. Therefore, step S1 specifically includes:

[0059] The lateral normal velocity increment is obtained by combining inertial measurement and control data. The lateral normal overload Ny and Nz are obtained according to the first formula. The lateral normal overload Ny and Nz include the lateral overload Nz and the normal overload Ny. The lateral normal velocity increment includes the lateral velocity increment and the normal velocity increment.

[0060] The first equation is:

[0061]

[0062]

[0063] In the formula, Ny represents the normal overload, and ΔW y1 G is the apparent velocity increment in the normal direction, and G0 is the equivalent gravitational acceleration, G0 = 9.80667 m / s². 2 Nz represents lateral overload, ΔW z1 The normal apparent velocity increment is t, which is a preset time, typically 0.01s, or 10ms.

[0064] It should be noted that in the rocket body coordinate system O1x1y1z1, the origin O1 is the rocket's center of mass. The O1x1 axis is the axis of symmetry of the rocket's outer shell, pointing towards the rocket's nose. The O1y1 axis lies within the rocket's principal plane of symmetry, which coincides with the xOy plane of the launch coordinate system at the instant of launch. The O1y1 axis is perpendicular to the O1x1 axis, and the O1x1 axis is perpendicular to the principal plane of symmetry. Viewed along the launch direction, the O1x1 axis points to the right. O1x1y1z1 is a right-handed rectangular coordinate system. The rocket body coordinate system can be simply referred to as the rocket system.

[0065] Step S2: Filter calculation. Filter the transverse normal overload.

[0066] In some embodiments, the transfer function of the filter used to filter transverse normal overload can be:

[0067]

[0068] Where T0 is the time constant, which is generally taken to be around 2.5.

[0069] It should be noted that in step S2, the transverse normal overloads Ny and Nz are filtered, and the filter needs to meet three requirements:

[0070] 1. Filter out the natural frequencies of the elastic arrow body (the natural frequencies of the arrow body may contain multiple orders or local frequencies);

[0071] 2. It can eliminate high-frequency noise in the system;

[0072] 3. Meet the requirement of smooth compensation command values.

[0073] Therefore, a low-pass filter is generally used, and a first-order inertial filter can be selected.

[0074] Step S3: Integrator Calculation. Integrate the filtered transverse normal overload to obtain the transverse normal velocity increment;

[0075] In some embodiments, in step S3, the discretized integrator, i.e., the integrator function for integrating the filtered transverse normal overload, can be:

[0076] C(i) = k·T·R(i) + C(i-1)

[0077] Where k is the integral gain, which can range from 2 to 3, T is the launch vehicle control cycle, which can be 10 ms, C(i) is the output of the current launch vehicle control cycle, C(i-1) is the output of the previous launch vehicle control cycle, and R(i) is the input of the current launch vehicle control cycle.

[0078] Step S4: Nonlinear processing. The transverse normal velocity increment is subjected to amplitude limiting and linear transition processing to obtain the compensation command value.

[0079] Specifically, when performing amplitude limiting, the amplitude can be selected based on the angle of attack generated by the wind field, generally with an angle of attack to rudder deflection ratio of 1:1. For better compensation, considering an angle of attack of α and a corresponding compensation command value amplitude limit of α, the compensation command value amplitude limit for limiting the lateral normal velocity increment can be:

[0080]

[0081] Where δ p Let dVy and dVz be the transverse normal velocity increments, and α be the angle of attack of the launch vehicle.

[0082] It should be noted that, from a control perspective, the lateral and normal directions are the same. That is, the limits for both the lateral and normal compensation command values ​​are defined by the formulas described above. However, when δ0 is the limit for the normal compensation command value, the corresponding δ... p δ is the normal velocity increment dVy; δ0 is the corresponding δ when the lateral compensation command value is limited. p dVz represents the lateral velocity increment.

[0083] When performing linear transition processing, considering that the need to introduce compensation command values ​​only arises at a certain altitude and with a certain dynamic pressure, and to avoid disturbances when introducing compensation command values, a flight altitude variable is introduced. The compensation amount is then linearly introduced into the control loop along with the altitude. Therefore, when performing linear transition processing on the lateral normal velocity increment, the compensation command value coefficient is:

[0084]

[0085] In the formula, h1 is the altitude, approximately 2500m; h2 is the preset altitude, approximately 20000m; and H is the flight altitude of the launch vehicle.

[0086] Therefore, the formula for compensating the instruction value δ2 is:

[0087] δ2=k b ·δ0

[0088] It should be noted that, since the rocket flies at an altitude between 5000m and 18000m and is subject to strong winds, the inventors of this application, through in-depth research and practice, set h1 to 2500m and h2 to 20000m, and finally obtained a better compensation command value coefficient.

[0089] Therefore, in some embodiments, step S4 specifically includes:

[0090] The launch vehicle's angle of attack and flight altitude are obtained through the attitude control system;

[0091] The compensation command value is obtained based on the transverse normal velocity increment and the second formula;

[0092] The second equation is:

[0093]

[0094]

[0095] δ2=k b ·δ0

[0096] In the formula, δ p α is the lateral normal velocity increment, h1 is the launch vehicle's angle of attack, h2 is the altitude, h is the preset altitude, H is the launch vehicle's flight altitude, and δ2 is the compensation command value.

[0097] Step S5: Rudder Command Calculation. Based on the sum of the rudder command value δ1 and the compensation command value δ2 obtained from the launch vehicle's attitude control system, the final rudder command value Δδ is obtained, i.e., Δδ = δ1 + δ2. The launch vehicle's actuators (e.g., servo motors) control the launch vehicle's attitude movement associated with the final rudder command value. Specifically, the servo motor rotates by a corresponding angle according to the final rudder command value, generating control force and control torque, which act on the rocket body, thereby manipulating the rocket body's attitude movement, enabling the rocket to fly with the wind, reducing the aerodynamic angle of attack (sideslip angle), and achieving the purpose of reducing aerodynamic loads.

[0098] The high-altitude wind load reduction control method for launch vehicles provided in this application has the following advantages:

[0099] 1. No wind field setup is required before launch, which can meet the mission-specific requirements of launch vehicles;

[0100] 2. It can reduce aerodynamic load, thereby reducing the structural strength requirements of the rocket body, reducing the weight of the rocket body, increasing carrying capacity, and resulting in high economic benefits.

[0101] 3. The control method is simple, has strong adaptability to changes in the upper-level wind field, and can still be applied when shear wind occurs, making it highly valuable for engineering applications.

[0102] Based on the same inventive concept, such as Figure 4 As shown in the embodiment of this application, a high-altitude wind load reduction control device for a launch vehicle is also provided, comprising:

[0103] The first calculation unit 10 is used to obtain the horizontal normal apparent velocity increment of the rocket body coordinate system and obtain the horizontal normal overload based on the horizontal normal apparent velocity increment.

[0104] Filtering unit 20 is used to filter transverse normal overload;

[0105] Integration unit 30 is used to integrate the filtered transverse normal overload to obtain the transverse normal velocity increment.

[0106] Processing unit 40 is used to perform amplitude limiting and linear transition processing on the transverse normal velocity increment to obtain the compensation command value;

[0107] The second calculation unit 50 is used to obtain the final rudder command value based on the sum of the rudder command value and the compensation command value obtained from the attitude control system of the launch vehicle; the actuator of the launch vehicle controls the launch vehicle to move in attitude associated with the final rudder command value.

[0108] In some embodiments, the first computing unit 10 is specifically used for:

[0109] The lateral normal velocity increment is obtained by combining inertial measurement units (INS) and the lateral normal overload is obtained according to the first formula. The lateral normal overload includes lateral overload and normal overload, and the lateral normal velocity increment includes lateral velocity increment and normal velocity increment.

[0110] The first equation is:

[0111]

[0112]

[0113] In the formula, Ny represents the normal overload, and ΔW y1 G0 is the normal apparent velocity increment, G0 is the equivalent gravitational acceleration, Nz is the lateral overload, and ΔW is the apparent velocity increment. z1 The value represents the lateral apparent velocity increment, and t represents the preset time, which can be 0.01s.

[0114] In some embodiments, the filtering unit 20 can be a filter, specifically used to filter the transverse normal overload through a transfer function, which can be:

[0115]

[0116] Where T0 is the time constant, and T0 can be 2.5.

[0117] In some embodiments, the integration unit 30 is specifically used to integrate the filtered transverse normal overload through a discretized integration element (i.e., an integration function), and the discretized integration element can be:

[0118] C(i) = k·T·R(i) + C(i-1)

[0119] Where k is the integral gain, T is the launch vehicle control cycle, C(i) is the output of the current launch vehicle control cycle, C(i-1) is the output of the previous launch vehicle control cycle, and R(i) is the input of the current launch vehicle control cycle. The range of k can be 2 to 3, and T can be 0.01s.

[0120] In some embodiments, the processing unit 40 is specifically used to obtain the angle of attack and flight altitude of the launch vehicle through the attitude control system, and to obtain the compensation command value according to the lateral normal velocity increment and the second formula.

[0121] The second equation is:

[0122]

[0123]

[0124] δ2=k b ·δ0

[0125] In the formula, δ p α is the lateral normal velocity increment, h1 is the launch vehicle's angle of attack, h2 is the altitude, h is the preset altitude, H is the launch vehicle's flight altitude, and δ2 is the compensation command value.

[0126] Specifically, h1 can be 2500m and h2 can be 20000m.

[0127] In summary, the high-altitude wind load reduction control method and device for launch vehicles provided in this application obtains the lateral normal overload based on the lateral normal apparent velocity increment output by the inertial measurement unit. After filtering and integration, the lateral normal velocity increment is obtained. After limiting and linear transition processing of the lateral normal velocity increment, a compensation command value is obtained. Based on the sum of the rudder command value obtained from the launch vehicle's attitude control system and the compensation command value, the final rudder command value is obtained. The final rudder command value is output to the launch vehicle's actuator (e.g., a servo motor). The actuator controls the launch vehicle to move at the attitude associated with the final rudder command value, thereby enabling the rocket body to fly with the wind, reducing the angle of attack (sideslip angle), and achieving the purpose of reducing aerodynamic load. The technical solution provided in this application, verified through mathematical simulation and semi-physical simulation, can control the rocket body load within 0.35g, solving the high-altitude wind load problem, further improving the adaptability to high-altitude wind field loads, reducing the structural weight of the rocket, increasing the carrying capacity, enhancing the rocket's de-tasking capability, and possessing extremely high engineering application value.

[0128] Other embodiments of this application will readily conceive of by those skilled in the art upon consideration of the specification and practice of the embodiments disclosed herein. This application is intended to cover any variations, uses, or adaptations of this application that follow the general principles of this application and include common knowledge or customary techniques in the art not disclosed herein. It should be understood that this application is not limited to the precise structures described above and shown in the accompanying drawings, and various modifications and changes can be made without departing from its scope. The scope of this application is limited only by the appended claims.

Claims

1. A method for controlling high-altitude wind load reduction in a launch vehicle, characterized in that, include: Obtain the lateral normal velocity increment of the rocket body coordinate system, and obtain the lateral normal overload based on the lateral normal velocity increment; Integrating the lateral normal overload, we obtain the lateral normal velocity increment; The transverse normal velocity increment is subjected to amplitude limiting and linear transition processing to obtain the compensation command value; The final rudder command value is obtained by summing the rudder command value obtained from the attitude control system of the launch vehicle and the compensation command value. The actuator of the launch vehicle controls the launch vehicle to move in attitude associated with the final rudder command value. The process of limiting and linearly transitioning the transverse normal velocity increment to obtain the compensation command value includes: The angle of attack and flight altitude of the launch vehicle are obtained through the attitude control system. The compensation command value is obtained based on the lateral normal velocity increment and the second formula; The second formula is: In the formula, For the transverse normal velocity increment, δ1 is the angle of attack of the launch vehicle, h1 is the altitude, h2 is the preset altitude, H is the flight altitude of the launch vehicle, and δ2 is the compensation command value.

2. The method for controlling high-altitude wind load reduction of a launch vehicle according to claim 1, characterized in that, The process of obtaining the lateral normal velocity increment of the rocket body coordinate system and obtaining the lateral normal overload based on the lateral normal velocity increment includes: The lateral normal velocity increment is obtained by combining inertial measurement and control systems, and the lateral normal overload is obtained according to the first formula; the lateral normal overload includes lateral overload and normal overload, and the lateral normal velocity increment includes lateral velocity increment and normal velocity increment. The first formula is: In the formula, Ny represents the normal overload. W y1 G0 is the normal apparent velocity increment, G0 is the equivalent gravitational acceleration, and Nz is the lateral overload. W z1 The lateral apparent velocity increment is given by t, which is the preset time.

3. The high-altitude wind load reduction control method for launch vehicles according to claim 2, characterized in that, The value of t is 0.01s.

4. The high-altitude wind load reduction control method for launch vehicles according to claim 1, characterized in that, Before the step of integrating the lateral normal overload to obtain the lateral normal velocity increment, the method further includes: The lateral normal overload is filtered.

5. The high-altitude wind load reduction control method for launch vehicles according to claim 4, characterized in that, The transfer function of the filter used to filter the transverse normal overload is: Where T0 is the time constant.

6. The high-altitude wind load reduction control method for launch vehicles according to claim 1, characterized in that, The integral function for integrating the transverse normal overload is: Where k is the integral gain and T is the launch vehicle control period. This is the output of the current launch vehicle control cycle. This is the output of the previous launch vehicle control cycle. This is the input for the current launch vehicle control cycle.

7. The high-altitude wind load reduction control method for launch vehicles according to claim 6, characterized in that, The range of k is 2 to 3, and the value of T is 0.01 s.

8. The method for controlling high-altitude wind load reduction of a launch vehicle according to claim 1, characterized in that, h1 is 2500m and h2 is 20000m.

9. A high-altitude wind load reduction control device for a launch vehicle, characterized in that, include: The first calculation unit is used to obtain the lateral normal velocity increment of the rocket body coordinate system and to obtain the lateral normal overload based on the lateral normal velocity increment. An integrator unit is used to integrate the lateral normal overload to obtain the lateral normal velocity increment. The processing unit is used to perform amplitude limiting and linear transition processing on the transverse normal velocity increment to obtain a compensation command value; The second calculation unit is used to obtain the final rudder command value based on the sum of the rudder command value obtained by the attitude control system of the launch vehicle and the compensation command value; the actuator of the launch vehicle controls the launch vehicle to move with the attitude associated with the final rudder command value. The processing unit is further configured to: The angle of attack and flight altitude of the launch vehicle are obtained through the attitude control system. The compensation command value is obtained based on the lateral normal velocity increment and the second formula; The second formula is: In the formula, For the transverse normal velocity increment, δ1 is the angle of attack of the launch vehicle, h1 is the altitude, h2 is the preset altitude, H is the flight altitude of the launch vehicle, and δ2 is the compensation command value.

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