A control method for a launch vehicle based on wind field characteristics

Abstract

The present application provides a control method for a launch vehicle based on wind field characteristics. The method includes: obtaining a wind field model generated by fitting through actually measured wind field data; determining the relative velocity of the launch vehicle relative to the atmosphere at the current moment in the aircraft coordinate system according to the velocity of the launch vehicle at the current moment in the launch coordinate system, the wind speed at the current moment, and a first conversion relationship; calculating an influence parameter through the relative velocity, and calculating the aerodynamic force at the current moment according to the influence parameter; combining the aerodynamic force, the preset initial value of the preset control parameter, and the flight program angle expression of the launch vehicle changing with time, integrating the differential equations of the centroid motion of the launch vehicle, and cyclically calculating the preset state variables at each moment; calculating the objective function of the launch vehicle according to the preset state variables at each moment; based on the sequential quadratic programming algorithm, calculating the objective functions corresponding to different preset control parameters through the objective function and the preset constraint conditions to determine the optimized value of the preset control parameter.

Description

Technical Field

[0001] The present application relates to the technical field of launch vehicles, and particularly to a control method for a launch vehicle based on wind field characteristics. Background Art

[0002] In traditional trajectory design, it is customary to take the velocity of the launch vehicle in the inertial coordinate system as the oncoming flow velocity of the standard atmosphere, and consider the movement of the atmosphere itself as an interference term. On the one hand, this method causes a certain deviation between the aerodynamic force and aerodynamic moment of the launch vehicle and those in the static atmosphere, thus interfering with the trajectory. On the other hand, the motion state of the launch vehicle obtained by the measuring equipment cannot reflect the real change, resulting in the stable control system being unable to respond accurately and making it difficult to control the missile to fly according to the predetermined trajectory plan. Especially for the rocket launch under a static base, the missile speed is relatively low in the initial stage, and the influence of the wind vector cannot be ignored, which to a certain extent affects the normal implementation of the launch mission. Summary of the Invention

[0003] In view of this, the purpose of the present application is to provide at least a control method for a launch vehicle based on wind field characteristics. By constructing a wind field model according to the actually measured wind field data, determining the relative velocity of the launch vehicle relative to the atmosphere in the aircraft coordinate system according to the wind speed simulated by the wind field model, and calculating the aerodynamic force through the relative velocity, and combining the aerodynamic force, the preset initial value of the preset control parameter, and the expression of the flight program angle of the launch vehicle changing with time, integrating the differential equations of the centroid motion of the launch vehicle to calculate the preset state variables of the launch vehicle at each moment, and outputting the objective function of the preset state variables. By using the sequential quadratic programming algorithm to combine the objective function and the preset constraint conditions to change the values of different preset control parameters, and correspondingly calculating the objective functions corresponding to different preset control parameters respectively, to determine the optimized values of the preset control parameters, which is convenient to control the launch vehicle to execute the flight mission according to the optimized values of the preset control parameters, solving the technical problem that the influence of the wind field on the launch vehicle is not considered in the prior art, and achieving the technical effect of improving the success probability of the launch vehicle executing the flight mission.

[0004] The present application mainly includes the following aspects:

[0005] In a first aspect, an embodiment of the present application provides a control method for a launch vehicle based on wind field characteristics, and the method includes: obtaining a wind field model generated by fitting through actually measured wind field data; determining the relative velocity of the launch vehicle relative to the atmosphere at the current moment in the aircraft coordinate system according to the velocity of the launch vehicle at the current moment in the launch coordinate system, the wind speed at the current moment simulated in the wind field model in the launch coordinate system, and a first conversion relationship, where the first conversion relationship is used to convert the launch coordinate system to the aircraft coordinate system; calculating an influence parameter through the relative velocity, and calculating the aerodynamic force at the current moment according to the influence parameter; combining the aerodynamic force, a preset initial value of a preset control parameter, and an expression of the flight program angle of the launch vehicle varying with time, integrating the differential equations of the centroid motion of the launch vehicle, calculating the preset state variables of the launch vehicle at the next moment, and obtaining the preset state variables at each moment in this way; calculating an objective function of the launch vehicle according to the preset state variables at each moment; based on the sequential quadratic programming algorithm, changing the value of the preset control parameter through the objective function and preset constraint conditions, and calculating the corresponding objective function under different preset control parameters to determine the optimized value of the preset control parameter.

[0006] Optionally, the aerodynamic force is calculated in the following manner: calculating an influence parameter and dynamic pressure that affect the aerodynamic coefficient through the velocities of the relative velocity under each coordinate axis of the aircraft coordinate system, where the influence parameter includes the angle of attack and the sideslip angle; calculating the aerodynamic force according to the angle of attack, the sideslip angle, the aerodynamic area of the launch vehicle, the dynamic pressure, and the aerodynamic coefficient selected through the influence parameter.

[0007] Optionally, the flight program angle expression is used to describe the program angles of each time period during the flight of the launch vehicle, and the flight program angle expression includes:

[0008]

[0009] where, is the flight program angle varying with time, 0≤t<t1 is the vertical flight section of the launch vehicle, θ is the ballistic inclination angle, α(t) is the angle of attack varying with time, t1≤t<t2 is the turning section of the launch vehicle, is the preset program angle, t2≤t<t3 is the constant flight program angle section of the launch vehicle, is the slope of the flight program angle that remains constant starting from time t2, t3≤t<t4 is the equal slope turning section of the launch vehicle, is the pitch program angle that remains constant before the shutdown point of the launch vehicle, and t≥t4 is the aiming section of the launch vehicle in the second-stage flight phase.

[0010] Optionally, the preset control parameters include the firing azimuth angle, the maximum value of the absolute value of the angle of attack within the turning section, the time when the angle of attack reaches the extreme value within the turning section, the slope of the flight program angle, and the starting time of the aiming section in the second-stage flight phase of the launch vehicle. Among them, the angle of attack varying with time is described by the following formula:

[0011] α(t) = -4×α M ×Z×(1 - Z)

[0012]

[0013] where α(t) is the angle of attack varying with time, α M is the maximum value of the absolute value of the angle of attack within the turning section, a is a preset constant value, and t m is the time when the angle of attack reaches the extreme value within the turning section.

[0014] Optionally, the differential equations of the centroid motion of the launch vehicle are described by the following formula:

[0015]

[0016] where is the acceleration, is the second transformation relationship, which is used to transform the aircraft coordinate system to the launch coordinate system, is the apparent acceleration in the aircraft coordinate system, ω e is the earth rotation speed, R0 is the coordinate vector of the earth center in the launch coordinate system, is the derivative of the position vector X with respect to time, V is the velocity of the launch vehicle at each moment in the launch coordinate system, and g is the acceleration due to gravity.

[0017] Optionally, the objective function of the launch vehicle is calculated in the following manner: Based on the preset state variables at each moment, the objective parameter values of the launch vehicle are calculated. The objective parameter values include at least one of the following: the payload under the given orbit injection conditions, the elliptical orbit under the given payload, and the range; Based on each objective parameter value and its corresponding preset weight factor, the objective function of the launch vehicle is calculated.

[0018] Optionally, the optimized value of the preset control parameter is determined as follows: Based on the sequential quadratic programming algorithm, determine the current search direction and the current search step corresponding to the objective function and the preset constraint condition; according to the current search direction and the current search step, update the preset control parameter to obtain the updated value of the preset control parameter; determine whether the preset convergence condition is satisfied. If the preset convergence condition is not satisfied, return to recalculate the preset control parameter of the launch vehicle until the preset convergence condition is satisfied, and use the updated value of the preset control parameter corresponding to when the convergence condition is reached as the optimized value of the preset control parameter.

[0019] In a second aspect, an embodiment of the present application further provides a launch vehicle control device based on wind field characteristics. The device includes: an acquisition module that acquires a wind field model generated by fitting through actually measured wind field data; a determination module that determines an expression of the relative velocity of the launch vehicle relative to the atmosphere at the current moment in the aircraft coordinate system according to the velocity of the launch vehicle at the current moment in the launch coordinate system, the wind speed at the current moment simulated in the wind field model in the launch coordinate system, and a first conversion relationship, where the first conversion relationship is used to convert the launch coordinate system to the aircraft coordinate system; a first calculation module that calculates an influence parameter through the relative velocity expression and calculates the aerodynamic force at the current moment according to the influence parameter; a second calculation module that integrates the differential equations of the centroid motion of the launch vehicle by combining the aerodynamic force, the preset initial value of the preset control parameter, and the expression of the flight program angle of the launch vehicle changing with time, and calculates the preset state variable of the launch vehicle at the next moment, and obtains the preset state variables at each moment in this way; a third calculation module that calculates the objective function of the launch vehicle according to the preset state variables at each moment; an optimization module that, based on the sequential quadratic programming algorithm, changes the value of the preset control parameter through the objective function and the preset constraint condition, and calculates the objective function corresponding to different preset control parameters to determine the optimized value of the preset control parameter.

[0020] In a third aspect, an embodiment of the present application further provides an electronic device, including: a processor, a memory, and a bus. The memory stores machine-readable instructions executable by the processor. When the electronic device runs, the processor communicates with the memory through the bus. When the machine-readable instructions are run by the processor, the steps of the method for controlling a launch vehicle based on wind field characteristics described in the first aspect or any possible implementation manner in the first aspect are executed.

[0021] Fourthly, an embodiment of the present application further provides a computer-readable storage medium, on which a computer program is stored. When the computer program is run by a processor, it executes the steps of the method for controlling a launch vehicle based on wind field characteristics described in the first aspect or any possible implementation manner in the first aspect.

[0022] An embodiment of the present application provides a method for controlling a launch vehicle based on wind field characteristics. The method includes: obtaining a wind field model generated by fitting through actually measured wind field data; determining the relative velocity of the launch vehicle relative to the atmosphere at the current moment in the aircraft coordinate system according to the velocity of the launch vehicle at the current moment in the launch coordinate system, the wind speed at the current moment simulated in the wind field model, and a first conversion relationship, where the first conversion relationship is used to convert the launch coordinate system to the aircraft coordinate system; calculating an influence parameter through the relative velocity, and calculating the aerodynamic force at the current moment according to the influence parameter; combining the aerodynamic force, the preset initial value of the preset control parameter, and the expression of the flight program angle of the launch vehicle changing with time, integrating the differential equations of the centroid motion of the launch vehicle to calculate the preset state variables of the launch vehicle at the next moment, and thus obtaining the preset state variables at each moment in a loop; calculating the objective function of the launch vehicle according to the preset state variables at each moment; based on the sequential quadratic programming algorithm, changing the values of the preset control parameters through the objective function and the preset constraint conditions, and calculating the corresponding objective functions under different preset control parameters to determine the optimized value of the preset control parameter. By constructing a wind field model according to the actually measured wind field data, determining the relative velocity of the launch vehicle relative to the atmosphere in the aircraft coordinate system according to the wind speed simulated by the wind field model, and calculating the aerodynamic force through the relative velocity, so as to integrate the differential equations of the centroid motion of the launch vehicle by combining the aerodynamic force, the preset initial value of the preset control parameter, and the expression of the flight program angle of the launch vehicle changing with time, calculate the preset state variables of the launch vehicle at each moment, and output the objective function of the preset state variables. By using the sequential quadratic programming algorithm to combine the objective function and the preset constraint conditions to change the values of different preset control parameters, and calculating the corresponding objective functions corresponding to different preset control parameters respectively to determine the optimized value of the preset control parameter, it is convenient to control the launch vehicle to execute the flight mission according to the optimized value of the preset control parameter, solves the technical problem that the influence of the wind field on the launch vehicle is not considered in the prior art, and achieves the technical effect of improving the success probability of the launch vehicle executing the flight mission.

[0023] To make the above objects, features, and advantages of the present application more obvious and understandable, the following specifically gives preferred embodiments and cooperates with the attached drawings for detailed description as follows. Description of the Drawings

[0024] To more clearly illustrate the technical solutions of the embodiments of the present application, the accompanying drawings required for the embodiments will be briefly introduced below. It should be understood that the following drawings only show some embodiments of the present application and should not be regarded as limiting the scope. For those of ordinary skill in the art, without creative efforts, other related drawings can also be obtained based on these drawings.

[0025] Figure 1 The flowchart of a carrier rocket control method based on wind field characteristics provided by the embodiments of the present application is shown.

[0026] Figure 2 The curve graph showing the change of wind speed with height in the wind field model provided by the embodiments of the present application is shown.

[0027] Figure 3 The curve showing the change of wind speed with height in the wind field model provided by the embodiments of the present application is shown.

[0028] Figure 4 The curve showing the change of wind direction with height in the wind field model provided by the embodiments of the present application is shown.

[0029] Figure 5 The design curve graph showing the program angles designed corresponding to the wind speed considering the wind field model and the wind speed without considering the wind field model respectively provided by the embodiments of the present application is shown.

[0030] Figure 6 The schematic diagram showing the maximum aerodynamic force corresponding to the wind field model considered and the maximum aerodynamic force corresponding to the wind field model not considered provided by the embodiments of the present application is shown.

[0031] Figure 7 The functional module diagram of a carrier rocket control device based on wind field characteristics provided by the embodiments of the present application is shown.

[0032] Figure 8 The structural schematic diagram of an electronic device provided by the embodiments of the present application is shown. Detailed implementation manners

[0033] To make the objectives, technical solutions, and advantages of the embodiments of this application clearer, the technical solutions in the embodiments of this application will be clearly and completely described below in conjunction with the accompanying drawings in the embodiments of this application. It should be understood that the accompanying drawings in this application are only for the purposes of illustration and description, and are not used to limit the protection scope of this application. Additionally, it should be understood that the schematic drawings are not drawn to actual scale. The flowcharts used in this application illustrate the operations implemented according to some embodiments of this application. It should be understood that the operations in the flowchart may not be implemented in sequence, and steps without a logical context relationship may be reversed or implemented simultaneously. In addition, those skilled in the art can add one or more other operations to the flowchart or remove one or more operations from the flowchart under the guidance of the content of this application.

[0034] Furthermore, the described embodiments are only a part of the embodiments of this application, rather than all of the embodiments. The components of the embodiments of this application usually described and illustrated in the accompanying drawings here can be arranged and designed in various different configurations. Therefore, the following detailed description of the embodiments of this application provided in the accompanying drawings is not intended to limit the scope of this application that is required to be protected, but only represents the selected embodiments of this application. Based on the embodiments of this application, all other embodiments obtained by those skilled in the art without creative efforts fall within the protection scope of this application.

[0035] In the prior art, trajectory design plays an extremely important role in the overall design of launch vehicles and runs through the entire process of overall design. The overall scheme, design parameters, operational performance, flight plan, etc. of the rocket are all closely related to the flight trajectory. Considering from the trajectory, both theoretical research and launch tests have shown that the deviation of atmospheric parameters from their standard values has a greater impact on the movement of launch vehicles. In traditional trajectory design, it is customary to take the velocity of the launch vehicle in the inertial coordinate system as the oncoming flow velocity of the standard atmosphere and consider the movement of the atmosphere itself as an interference term. On the one hand, this method causes a certain deviation between the aerodynamic force and aerodynamic moment of the launch vehicle and those in the stationary atmosphere, thus interfering with the trajectory. On the other hand, the motion state of the launch vehicle obtained by the measuring equipment cannot reflect the real changes, resulting in the stable control system being unable to respond accurately and making it difficult to control the missile to fly according to the predetermined trajectory plan. Especially for the rocket launch under a static base, the missile speed is relatively low in the initial stage, and the influence of the wind vector cannot be ignored, which to a certain extent affects the normal implementation of the launch mission.

[0036] Based on this, the embodiments of the present application provide a control method for a launch vehicle based on wind field characteristics. By using a wind field model constructed according to the actually measured wind field data, the relative velocity of the launch vehicle relative to the atmosphere in the aircraft coordinate system is determined according to the wind speed simulated by the wind field model, and the aerodynamic force is calculated through the relative velocity. Then, by combining the aerodynamic force, the preset initial value of the preset control parameter, and the flight program angle expression of the launch vehicle changing with time, the differential equations of the centroid motion of the launch vehicle are integrated to calculate the preset state variables at each moment, and the objective function of the preset state variables is output. The sequential quadratic programming algorithm is used to change the values of different preset control parameters in combination with the objective function and the preset constraint conditions, and the objective functions corresponding to different preset control parameters are calculated respectively to determine the optimized values of the preset control parameters, so as to control the launch vehicle to execute the flight mission according to the optimized values of the preset control parameters, solving the technical problem that the influence of the wind field on the launch vehicle is not considered in the prior art, and achieving the technical effect of improving the success probability of the launch vehicle executing the flight mission. Specifically as follows:

[0037] Please refer to Figure 1 , Figure 1 which is a flowchart of a control method for a launch vehicle based on wind field characteristics provided by the embodiments of the present application. As Figure 1 shown, the control method for a launch vehicle based on wind field characteristics provided by the embodiments of the present application includes the following steps:

[0038] S101: Obtain a wind field model generated by fitting the actually measured wind field data.

[0039] Among them, the wind field data includes wind speed and wind direction. Although the forms of the atmospheric wind field are various and diverse, and will change continuously with factors such as terrain, longitude and latitude, altitude, temperature, air density, time, etc. However, through long-term research on the wind field at a fixed location, it is found that the change of the wind field shows obvious statistical characteristics. That is, according to the motion effect of the wind on the launch vehicle, the wind speed can be regarded as three components with different characteristics, including steady wind, shear wind, and gust wind. Steady wind refers to the horizontal wind that changes very slowly relative to time, and the curve of its change with height in a certain place and season can be obtained by statistically analyzing a large amount of observation data. Shear wind is the horizontal wind that only appears at a certain height, and its wind speed increases rapidly with the increase of height and then decreases rapidly. The time for the launch vehicle to pass through the shear wind area is very short, generally within 2s to 3s, so it is only affected by the short-term wind force, just like an impulse torque. Gust wind refers to the local disturbed air flow in the air around the launch vehicle, which can usually be regarded as a stationary random process and is described by the spectral density.

[0040] Exemplarily, please refer to Figure 2 , Figure 2 which is a graph showing the change of wind speed with height in the wind field model provided by the embodiments of the present application. AsFigure 2 As shown, for the curve of wind speed varying with height in the wind field model provided by the embodiments of the present application, the abscissa is the wind speed with the unit of meters per second, and the ordinate is the height with the unit of kilometers. Among them, a is the wind speed curve, b is the steady wind speed curve, c is the maximum wind speed envelope curve, and d is the shear wind curve. That is to say, the wind speed curve a can be understood as being obtained by fusing the steady wind speed curve b, the maximum wind speed envelope curve c, and the shear wind curve d. The wind speed curve a may also include gusts. Among them, the variation rates of different wind speed curves with height are different. The steady wind speed curve has the longest variation period with height, the shear wind comes second, and the gust is the shortest.

[0041] That is to say, the wind field at each flight time can be obtained by linearly superimposing gusts and the maximum steady wind speed, and the shear wind is only considered in the most severe cases. Furthermore, a fixed location is selected as the launch site of the launch vehicle, and the wind speed and wind direction at the launch site of the launch vehicle are detected for a long time, and a wind field model with different wind speed and wind direction changes over time is fitted through the detected wind speed and wind direction, and the wind speed and wind direction at each time are simulated with the wind field model. Moreover, the wind field model shows a periodic change in a statistical sense with the seasons. Exemplarily, by collecting the wind speed and wind direction in a fixed month in recent years, a wind field model reflecting the wind field characteristics of the fixed month is obtained, so that the wind field model can simulate the wind speed and wind direction of the fixed month.

[0042] S102: Determine the expression of the relative velocity of the launch vehicle relative to the atmosphere at the current moment in the aircraft coordinate system according to the velocity of the launch vehicle at the current moment in the launch coordinate system, the wind speed simulated in the wind field model at the current moment in the launch coordinate system, and the first conversion relationship.

[0043] Among them, the first conversion relationship is used to convert the launch coordinate system to the aircraft coordinate system. The launch coordinate system in the present application refers to a coordinate system with the origin O at the launch point, the OX axis pointing in the launch aiming direction in the horizontal plane of the launch point, the OY axis along the plumb line of the launch point upward, and the OZ axis forming a right-handed rectangular coordinate system with the OX axis and the OY axis. The aircraft coordinate system is also called the body coordinate system or the rocket body coordinate system, with the origin O at the center of mass of the aircraft, the OX axis along the longitudinal axis of the aircraft pointing to the head; the OY axis is perpendicular to the OX axis in the longitudinal symmetry plane of the aircraft and points upward; the OZ axis forms a right-handed rectangular coordinate system with the OX axis and the OY axis.

[0044] Among them, the aircraft refers to the launch vehicle in the present application. Furthermore, when the flight mission of the launch vehicle is known, the launch coordinate system can be constructed based on the known launch aiming direction of the launch vehicle, and the aircraft coordinate system can be constructed based on the center of mass, longitudinal axis, and longitudinal symmetry plane of the launch vehicle, and the first conversion relationship for converting the launch coordinate system to the aircraft coordinate system can be obtained.

[0045] Exemplarily, the wind vector at a certain moment in the wind field model is represented in the launch coordinate system as:

[0046]

[0047] In formula (1), \(W\) is the wind vector at each moment in the launch coordinate system, is the transformation matrix from the north-east-down coordinate system to the launch coordinate system, \(W\) E is the wind vector at each moment of the launch site of the launch vehicle in the north-east-down coordinate system in the wind field model, \(a_0\) is the wind direction, \(a\) c is the wind speed constant. Among them, the origin of the north-east-down coordinate system can be the same as the origin of the launch coordinate system, or a random origin can be selected. The \(X\)-axis points north, the \(Y\)-axis points up, and the \(Z\)-axis points east.

[0048] That is to say, the wind field model can directly simulate the wind vector at each moment in the launch coordinate system, or simulate the wind vector at each moment in the north-east-down coordinate system and then calculate the wind vector at a certain moment in the launch coordinate system through the transformation matrix from the north-east-down coordinate system to the launch coordinate system. The present application does not limit the coordinate system used in the wind field model, and only needs to determine the wind vector in the launch coordinate system through the transformation matrix.

[0049] Furthermore, when considering the wind vector, the launch vehicle represents its relative velocity to the atmosphere in the launch coordinate system through the following formula:

[0050] \(V\) W = \(V - W\) (2)

[0051] In formula (2), \(V\) W is the relative velocity of the launch vehicle to the atmosphere in the launch coordinate system at each moment, \(V\) is the velocity of the launch vehicle in the launch coordinate system at each moment, and \(W\) is the wind vector at each moment in the launch coordinate system.

[0052] That is to say, by calculating the velocity of the launch vehicle at the current moment in the launch coordinate system in step S104 and combining it with the wind vector at the current moment in the wind field model, the relative velocity of the launch vehicle to the atmosphere at the current moment in the launch coordinate system can be calculated. Thus, through the relative velocity to the atmosphere at the current moment, the subsequent steps are executed until the differential equations of the centroid motion of the launch vehicle are integrated, and the velocity of the launch vehicle at the next moment in the launch coordinate system is solved. Then, step S102 is executed again to achieve a cycle and calculate the preset state variables of the launch vehicle at each time. And the initial velocity of the launch vehicle in the launch coordinate system is known.

[0053] Therefore, through the transformation matrix from the launch coordinate system to the vehicle coordinate system and the velocity of the launch vehicle relative to the atmosphere in the launch coordinate system, the expression for the relative velocity of the launch vehicle relative to the atmosphere in the vehicle coordinate system can be determined:

[0054]

[0055] In Formulas (3) and (4), V1 is the relative velocity of the launch vehicle relative to the atmosphere in the vehicle coordinate system at each moment, V x1 is the relative velocity of the launch vehicle relative to the atmosphere along the X-axis in the vehicle coordinate system at each moment, V y1 is the relative velocity of the launch vehicle relative to the atmosphere along the Y-axis in the vehicle coordinate system at each moment, V z1 is the relative velocity of the launch vehicle relative to the atmosphere along the Z-axis in the vehicle coordinate system at each moment, is the transformation matrix from the launch coordinate system to the vehicle coordinate system, V W is the velocity of the launch vehicle relative to the atmosphere in the launch coordinate system at each moment, ψ is the yaw program angle, is the pitch program angle, γ is the roll program angle. Generally, only is designed, and the other two program angles can be ignored. In this application, the pitch program angle is subsequently represented by .

[0056] That is to say, in this application, by considering the influence of the wind field on the launch vehicle, the velocity of the launch vehicle relative to the atmosphere and the aerodynamic force are calculated to achieve passive load reduction.

[0057] S103: Calculate the influence parameters through the relative velocity, and calculate the aerodynamic force at the current moment based on the influence parameters.

[0058] When the launch vehicle moves relative to the atmosphere, how to determine the aerodynamic force acting on the rocket is a rather complex problem and it is difficult to accurately determine through theoretical calculation. Currently, a method combining calculation using aerodynamics theory and calibration through aerodynamic experiments is used. The aerodynamic experiment is carried out in a wind tunnel that can generate a uniform airflow with a certain Mach number. Exemplarily, a scaled-down physical model of the launch vehicle is obtained, and the physical model is placed in the wind tunnel for a wind tunnel test, so that the airflow blows over this model at a certain Mach number. At this time, the model is placed in the wind tunnel and has no power, and only the velocity V W of the launch vehicle relative to the atmosphere in the launch coordinate system is used for simulation, and various V WThe corresponding aerodynamic coefficients under multiple sideslip angle ranges and multiple angle of attack ranges. The aerodynamic coefficients are calculated by measuring the aerodynamic forces acting on this model, and then applying the similarity transformation principle to obtain the aerodynamic forces experienced by the physical object at these Mach numbers. Among them, the aerodynamic forces are divided into lift, drag, and side force.

[0059] During the rocket development process, specialized personnel in aerodynamics, based on the rocket's shape, use the methods mentioned above to provide the necessary tables, curves, and mathematical models for calculating the aerodynamic forces of this type of launch vehicle to facilitate the selection of aerodynamic coefficients.

[0060] Specifically, the aerodynamic forces are calculated in the following way: The influence parameters and dynamic pressure are calculated through the velocities of the relative velocity in each coordinate axis of the vehicle coordinate system. The influence parameters include the angle of attack and the sideslip angle; the aerodynamic forces are calculated based on the angle of attack, the sideslip angle, the aerodynamic area of the launch vehicle, the dynamic pressure, and the aerodynamic coefficient selected through the influence parameters.

[0061] Among them, the influence parameters are calculated through the following formula:

[0062]

[0063] In Formulas (5) and (6), α is the angle of attack, β is the sideslip angle, V x1 is the relative velocity of the launch vehicle relative to the atmosphere along the X-axis of the vehicle coordinate system at each moment, V y1 is the relative velocity of the launch vehicle relative to the atmosphere along the Y-axis of the vehicle coordinate system at each moment, V z1 is the relative velocity of the launch vehicle relative to the atmosphere along the Z-axis of the vehicle coordinate system at each moment.

[0064] That is to say, the relative velocity of the launch vehicle relative to the atmosphere in the vehicle coordinate system at each moment includes the relative velocity of the X-axis of the vehicle coordinate system relative to the atmosphere, the relative velocity of the Y-axis of the vehicle coordinate system relative to the atmosphere, and the relative velocity of the Z-axis of the vehicle coordinate system relative to the atmosphere.

[0065] Among them, the dynamic pressure is calculated through the following formula:

[0066]

[0067] In Formula (7), q is the dynamic pressure, ρ is the air density, V h is the resultant velocity of the launch vehicle in the vehicle coordinate system, which can be obtained by summing the squares of V x1 、V y1 、V z1 respectively and then taking the square root.

[0068] Among them, the aerodynamic force is calculated by the following formula:

[0069]

[0070] In formula (8), A is the drag force, N is the lift force, Z1 is the side force, C A is the aerodynamic coefficient corresponding to the drag force, q is the dynamic pressure, S M is the aerodynamic area of the launch vehicle, α is the angle of attack, β is the sideslip angle, is the aerodynamic coefficient corresponding to the lift force, is the aerodynamic coefficient corresponding to the side force. The aerodynamic coefficient is the aerodynamic coefficient corresponding to the angle of attack and sideslip angle within their respective ranges determined through wind tunnel tests.

[0071] That is to say, the influencing parameters are used to influence the selection of the aerodynamic coefficient, and the influencing parameters can also include parameters such as altitude and Mach. At the current time, by determining the velocities of the launch vehicle along each coordinate axis in the aircraft coordinate system, the influencing factors such as the angle of attack and sideslip angle of the launch vehicle at the current time can be calculated, so as to select the appropriate aerodynamic coefficient through the data table, curve or mathematical model under the constraints of different variables such as altitude, Mach, angle of attack, and sideslip angle obtained in advance, and calculate the dynamic pressure through the velocities of the launch vehicle along each coordinate axis in the aircraft coordinate system, thereby calculating the aerodynamic force of the launch vehicle through formula (8).

[0072] That is to say, the angle of attack, sideslip angle and dynamic pressure are calculated through the relative velocity of the launch vehicle relative to the atmosphere at the current moment in the aircraft coordinate system, and then through the aerodynamic coefficients corresponding to the angle of attack, sideslip angle and their respective ranges, thereby calculating the aerodynamic force at the current moment through the angle of attack, sideslip angle, dynamic pressure and aerodynamic coefficient. Furthermore, when the relative velocity of the launch vehicle relative to the atmosphere in the aircraft coordinate system at each moment is known, the aerodynamic force can be calculated in this way.

[0073] S104: Integrate the differential equations of the centroid motion of the launch vehicle by combining the aerodynamic force, the preset initial value of the preset control parameter, and the flight program angle expression of the launch vehicle varying with time, calculate the preset state variables of the launch vehicle at the next moment, and thus obtain the preset state variables at each moment by cycling.

[0074] Among them, the flight program angle expression is used to describe the program angles of each time period during the flight of the launch vehicle. The preset initial value of the preset control parameter can be set by technicians by combining experimental data, experience, and formulas, etc.

[0075] Among them, the selection of the flight program angle mainly refers to the design of the flight program angle in the active section. This is because the passive section trajectory depends on the motion parameters at the shutdown point (separation of the satellite and the rocket) in the active section. Once the motion parameters at the shutdown point are determined, the trajectory shape and range in the passive section are also determined. Usually, to reduce the aerodynamic load acting on the launch vehicle in the atmosphere during the flight section, which causes a decrease in the flight speed of the launch vehicle and further results in aerodynamic drag loss and angle of attack loss. Exemplarily, the launch vehicle of the present application is a two-stage launch vehicle. A gravity-turn trajectory is adopted on the flight trajectory in the first-stage flight phase, that is, shortly after the rocket takes off vertically and each subsystem works to reach a stable state, a programmed turn is implemented; before the flight speed approaches the transonic section, the rocket is controlled to fly at a zero angle of attack until the engine in the first-stage flight phase is close to shutdown.

[0076] Furthermore, when the engine in the first-stage flight phase shuts down, generally speaking, the rocket has basically flown out of the dense atmosphere. At this time, the aerodynamic force on the rocket body is very small. Then, according to the force conditions during its flight, when designing the flight trajectory in the second-stage flight phase, the influence of the aerodynamic load on the trajectory can be ignored, and mainly consider making the payload reach the predetermined re-entry state. To improve the stability of the rocket, before and after the engine starts, shuts down, and stage separation, it is required that the rocket maintains a constant program angle during flight and tries to reduce the flight angle of attack as much as possible. In addition, the designed variation law of the program pitch angle should be as simple as possible to facilitate the implementation of the control and guidance system.

[0077] Specifically, the flight program angle expression includes:

[0078]

[0079] Among them, in formula (9), is the flight program angle that changes with time. 0≤t<t1 is the vertical flight section of the launch vehicle, θ is the ballistic inclination angle, α(t) is the angle of attack that changes with time, t1≤t<t2 is the turning section of the launch vehicle, is the preset program angle, t2≤t<t3 is the constant flight program angle section of the launch vehicle, is the flight program angle slope that remains constant starting from time t2, t3≤t<t4 is the constant slope turning section of the launch vehicle, is the constant pitch program angle maintained by the launch vehicle before the shutdown point, and t≥t4 is the aiming section of the launch vehicle in the second-stage flight phase.

[0080] Among them, launch vehicles generally choose vertical takeoff to overcome the disadvantages of inclined launch, simplify the launch equipment, only requiring a launch pad with a simple structure, and at the same time, the takeoff moment of the launch vehicle can be kept stable. The angle of attack in the vertical flight section 0 ≤ t < t1 of the launch vehicle is zero, and the ballistic inclination angle is equal to the program angle. The selection of t1 should be reasonable. If the time of the vertical flight section is too long, it will increase the gravity loss of speed and require a large normal force due to excessive speed during turning; if the time of the vertical flight section is too short, it is very likely that the engine has not reached the rated working state, and the actuators of the control system cannot generate a large enough control force, thus affecting the ballistic performance. Therefore, usually, the time of the vertical flight section needs to be maintained at least until the moment when the engine enters the rated working state, and at this time, the control mechanism can also normally control the turning. During preliminary design, the end moment t1 of the vertical flight section can be determined according to the empirical relationship curve between the vertical ascent time and the thrust-to-weight ratio ν0 of the launch vehicle. Exemplarily, the end moment is calculated by the following formula:

[0081]

[0082] Among them, in formula (10), t1 is the end moment of the vertical flight section, ν0 is the thrust-to-weight ratio of the launch vehicle, and the thrust-to-weight ratio refers to the ratio of the thrust generated by the engine of the launch vehicle to the mass of the launch vehicle. The thrust here can be considered a preset known quantity.

[0083] Among them, in the turning section t1 ≤ t < t2 of the launch vehicle, it can be further divided according to whether the angle of attack is zero. In the early stage of the turning section (t1 ≤ t < t ′ ), the angle of attack is not zero, and the launch vehicle realizes a turning with an angle of attack. According to the requirements, it should end before the aerodynamic force changes sharply to transonic speed to reduce the aerodynamic load and aerodynamic interference. When the Mach number M(t ′ ) is 0.7 to 0.8, the angle of attack shrinks to zero. Furthermore, in the later stage of the turning section (t ′ ≤ t < t2), the angle of attack is zero, and it turns slowly only relying on the normal component of gravity, that is, gravity turn, and t2 is used to indicate the end moment of the turning section of the launch vehicle, or the shutdown time of the minimum range of the medium and short-range rocket.

[0084] Among them, the angle of attack changing with time is described by the following formula:

[0085] a(t) = -4 × α M × Z × (1 - Z) (11)

[0086]

[0087] Among them, in formulas (11) to (13), α(t) is the angle of attack changing with time in the turning section of the launch vehicle, αM is the maximum value of the absolute value of the angle of attack within the turning section, α is a preset constant value, and t m is the time when the angle of attack reaches an extreme value within the turning section. From formula (11), it can be obtained that α(t) starts to rapidly reach a negative value, and then its absolute value starts to decrease, tending to zero at an exponential rate. The rate of tending to zero is determined by the parameter a. By taking the derivative of α(t), when the slope of α(t) approaches 0, that is, obtain Substituting into formula (12), the preset constant value can be obtained t m can be understood as the time corresponding to the slope of α(t) approaching 0.

[0088] Among them, the constant flight section t2 ≤ t < t3 of the launch vehicle can be divided into the aiming section of the first-stage flight phase and the initial flight section of the second-stage flight phase. A preset program angle is set in the initial flight section of the second-stage flight phase The purpose is to facilitate the separation of the first and second stages of the launch vehicle and reduce the influence of interference during separation on the movement of the missile. At the same time, it enables the engine in the second-stage flight phase to reach the rated operating state.

[0089] The purpose of setting in the equal-slope turning section t3 ≤ t < t4 of the launch vehicle is to reach the preset pitch program angle at the shutdown point of the active section, and the shutdown point is after the time point t4. The launch vehicle maintains a constant pitch program angle in the aiming section t ≥ t4 of the second-stage flight phase to improve the rocket stability.

[0090] Among them, the preset control parameters include the firing azimuth angle, the maximum value of the absolute value of the angle of attack within the turning section, the time when the angle of attack reaches an extreme value within the turning section, the slope of the flight program angle, and the starting time of the aiming section of the launch vehicle in the second-stage flight phase. Furthermore, the present application optimizes the values of the above preset control parameters to achieve that at least one target parameter value designed by the launch vehicle reaches an extreme value.

[0091] Specifically, the differential equations of the centroid motion of the launch vehicle are described by the following formula:

[0092]

[0093] Among them, in formula (14), is the acceleration, is the second transformation relationship, and the second transformation relationship is used to convert the aircraft coordinate system to the launch coordinate system, is the apparent acceleration in the aircraft coordinate system, that is, the acceleration generated by external forces other than gravity, which can be understood as the acceleration generated by the thrust, aerodynamic force, etc. of the aircraft, ω e is the earth's rotation speed, R0 is the coordinate vector of the earth's center in the launch coordinate system, is the derivative of the position vector X with respect to time, V is the velocity of the launch vehicle at each moment in the launch coordinate system, and g is the acceleration due to gravity.

[0094] That is to say, by selecting the program angle at the current moment from the preset initial values of the preset control parameters, the calculated aerodynamic force, and the flight program angle expression, and using the fourth-order Runge-Kutta method to integrate the differential equations of the center-of-mass motion, the preset state variables of the launch vehicle at the next moment are calculated. The preset state variables include the coordinates of the launch vehicle at the next moment in the aircraft coordinate system and the velocities of each coordinate axis.

[0095] Among them, the fourth-order Runge-Kutta method calculates the value of the next point by using the weighted average of four slopes within the time step, so as to calculate the coordinates and velocities of the launch vehicle in the aircraft coordinate system at each moment. The time step refers to the time difference between the current moment and the next moment.

[0096] That is to say, under the condition of the preset initial values of the preset control parameters, the coordinates and velocities at the next moment are solved through the aerodynamic force and program angle at the current moment, and the coordinates and velocities at the next moment are used as the new current moment for cycling, so as to solve the coordinates and velocities corresponding to each moment.

[0097] S105: Calculate the objective function of the launch vehicle according to the preset state variables at each moment.

[0098] Specifically, the objective function of the launch vehicle is calculated in the following way: according to the preset state variables at each moment, the objective parameter values of the launch vehicle are calculated, and the objective parameter values include at least one of the following: the payload under the given orbit injection conditions, the elliptical orbit under the given payload, and the range; according to each objective parameter value and its corresponding preset weight factor, the objective function of the launch vehicle is calculated.

[0099] That is to say, the trajectory optimization corresponds to one (single-objective optimization) or several (multi-objective optimization) objective parameter values, so as to tap the potential of the launch vehicle through trajectory optimization and improve its utilization value. At least one objective parameter value can be calculated through the velocity and position obtained by integration.

[0100] Exemplarily, the given orbit injection conditions generally correspond to preset values of the orbital elements, the orbit injection altitude, and the orbit injection velocity at the orbit injection moment. Only by changing the mass of the launch vehicle under different payloads, the mass m of the launch vehicle gIt is the sum of the payload mass and the structural mass. Based on this, the preset state variables at each moment under different launch vehicle masses are calculated, so that the launch vehicle mass that meets the given orbit injection conditions can be determined, and thus the payload under the given orbit injection conditions corresponding to the preset initial value of the preset control parameter can be obtained. Among them, the launch vehicle mass is related to the apparent acceleration calculated in the vehicle coordinate system.

[0101] Exemplarily, the following method is used to calculate the elliptical orbit and range under a given payload: Since the payload and structural mass of the launch vehicle are given, and the coordinates and velocities of the launch vehicle in the launch coordinate system at each moment can be calculated through formula (14), thus, the elliptical orbit and range can be obtained. The elliptical orbit can also be reflected by the semi-major axis and eccentricity of the orbit. The range refers to the range at the final position.

[0102] Furthermore, according to the calculation methods of each target parameter value, through the coordinates and velocities at each moment corresponding to the preset initial value of the preset control parameter, each target parameter value is calculated. Then, according to the preset weight factors of each target parameter value set in advance, by summing the products of each target parameter value and the preset weight factor, the value of the objective function of the launch vehicle is calculated. This objective function value represents the objective function value corresponding to the preset initial value of the preset control parameter, and this objective function value can be used to determine the preset iteration conditions of the subsequent sequential quadratic programming algorithm.

[0103] S106: Based on the sequential quadratic programming algorithm, the values of the preset control parameter are changed through the objective function and the preset constraint conditions, and the objective functions corresponding to different preset control parameters are calculated to determine the optimized value of the preset control parameter.

[0104] Specifically, the optimized value of the preset control parameter is determined in the following way: Based on the sequential quadratic programming algorithm, the current search direction and the current search step corresponding to the objective function and the preset constraint conditions are determined; according to the current search direction and the current search step, the preset control parameter is updated to obtain the updated value of the preset control parameter; it is determined whether the preset convergence condition is met. If the preset convergence condition is not met, the preset control parameter of the launch vehicle is recalculated until the preset convergence condition is met, and the updated value of the preset control parameter corresponding to when the convergence condition is reached is used as the optimized value of the preset control parameter.

[0105] Among them, the preset constraint conditions include the landing point constraint, the preset control parameter constraint, the dynamic pressure constraint, and the normal overload constraint. Among them, the landing point constraint refers to the constraint of the terminal ballistic parameters of the launch vehicle, such as the orbital element constraint of the injection point of the launch vehicle, etc. The preset control parameter constraint is used to describe the value range of each preset control parameter. Exemplarily, if the maximum value of the absolute value of the angle of attack in the turning section is too large, the turning amplitude is large, the terminal height is small and the speed is large; if the maximum value of the absolute value of the angle of attack in the turning section is small, the turning amplitude is small, the terminal height is large and the speed is small. Therefore, the optimized value of each preset control parameter should belong to its corresponding value range. From the perspective of the heat protection system design, the dynamic pressure should be within a certain range to ensure that the structure of the surface insulation material is not damaged. The control hinge moment increases with the increase of the dynamic pressure, and the dynamic pressure should also be kept not exceeding the dynamic pressure allowed by the maximum hinge moment required for controlling the control surface. At the same time, restricting the dynamic pressure can also ensure the stability of the launch vehicle during lateral flight to a certain extent. Exemplarily, the dynamic pressure should be less than or equal to the preset dynamic pressure value. The normal overload constraint means that the normal overload should be within a certain range. The maximum value of the normal overload mainly depends on the structural strength of the launch vehicle and the bearing range of the on-board equipment. To meet the structural design requirements, the normal overload is calculated by the following formula:

[0106]

[0107] In formula (15), |n y (t)| is the absolute value of the normal overload at each moment, P is the thrust of the launch vehicle, α is the angle of attack, Y is the component of the aerodynamic force in the direction perpendicular to the flight speed (i.e., the normal direction), m is the mass of the launch vehicle, and g is the acceleration due to gravity.

[0108] Among them, the preset constraint conditions can be changed according to different design tasks of the launch vehicle. For example, it is necessary to have the constraints on the debris fall zones of each stage, the height constraint at the shutdown point during flight tests, and the aerodynamic heat constraint, etc. Furthermore, the preset constraint conditions can be expressed as equality constraints and inequality constraints. The landing point constraint is an equality constraint, and the preset control parameter constraint, the dynamic pressure constraint, and the normal overload constraint are inequality constraints.

[0109] That is to say, a quadratic programming sub-problem of the sequential quadratic programming algorithm is constructed according to the objective function and the preset constraint conditions corresponding to the preset initial value of the preset control parameter. The search direction and step size are searched by solving the quadratic programming sub-problem, and then the value of the preset control parameter is updated by the search direction and compensation to obtain the updated value of the preset control parameter, and the updated value can satisfy the preset constraint conditions. And the coordinates and speeds at each moment corresponding to the updated value are calculated cyclically in the manner of steps S102 to S104 by using the updated value of the preset control parameter, and the objective function value is calculated when executing step S105. And it is judged whether the preset convergence condition is satisfied. The preset convergence condition means that the difference between the updated values of the preset control parameters calculated twice conforms to the preset difference range. If the preset convergence condition is satisfied, the optimized value of the preset control parameter can be directly obtained. Or after the preset convergence condition is satisfied, the updated value corresponding to the maximum payload under the given orbit injection condition, the maximum elliptical orbit under the given payload, and the maximum range among the updated values can be used as the optimized value of the preset control parameter. If the preset convergence condition is not satisfied, the updated value of the preset control parameter is modified to the preset initial value, the search direction and search step size are recalculated to determine the updated value of the preset control parameter, and the centroid motion differential equation set is re-integrated by using the fourth-order Runge-Kutta method to obtain the preset state variable, and the loop calculation is performed accordingly.

[0110] Exemplarily, in order to specifically analyze the design effect of the passive load reduction trajectory, first, a wind field modeling is constructed according to the historical wind field data of the launch site of the aircraft. Please refer to Figure 3 and Figure 4 , Figure 3 is the curve of the wind speed varying with height of the wind field model provided by the embodiment of the present application. The abscissa is the wind speed, the unit is meters per second, and the ordinate is the height, the unit is kilometers. Figure 4 is the curve of the wind direction varying with height of the wind field model provided by the embodiment of the present application. The abscissa is the wind direction angle, the unit is degrees, and the ordinate is the height, the unit is kilometers. As Figure 3 and Figure 4 shown, the wind direction angle is concentrated between 260° and 280°, the wind direction is westward, and the steady wind speed does not exceed 40 m / s.

[0111] Since, considering part of the influence of the wind field model, exemplarily, 30% of the wind speed influence is considered, that is, in order to reduce the error of the wind field model, only 30% of the wind speed can be used as the wind speed for actual calculation. Please refer to Figure 5 , Figure 5 is the design curve graph of the program angles respectively designed corresponding to the wind speed considering the wind field model and the wind speed not considering the wind field model provided by the embodiment of the present application. The abscissa is used to describe the flight time of the launch vehicle, and the ordinate is used to describe the program angle. As Figure 5As shown, the design of the program angle has not changed significantly. On this basis, simulations are carried out by considering the wind speed of the wind field model and the wind speed without considering the wind field model respectively to simulate the maximum aerodynamic force of the launch vehicle flight in the two cases. Please refer to Figure 6 , Figure 6 is a schematic diagram of the maximum aerodynamic force corresponding to the wind field model considered in the embodiment of the present application and the maximum aerodynamic force corresponding to the wind field model not considered. As Figure 6 shown, the maximum aerodynamic force decreases significantly when considering the wind field model to achieve passive load reduction. It reflects that under the same wind field action, the load reduction trajectory can enable the rocket body structure to withstand the influence of greater high-altitude wind interference, greatly improving the launch environment adaptability of the launch vehicle and increasing the launch probability.

[0112] Based on the wind field parameters of the existing launch site of the launch vehicle, the present application constructs a wind field model to reflect the wind field data under statistical characteristics. With the aim of increasing the launch probability and combining the actual mission requirements of the launch vehicle, the motion equation under the influence of the wind field model is given. Combining multiple preset constraints and the planned program angle, the theory of the numerical solution method for ballistic optimization is studied, that is, the initial values of the control parameters are given based on practical experience, and the sequential quadratic programming method is used to solve the nonlinear equation with constraints to obtain the ballistic parameters that meet the requirements. The simulation results show that the method in this paper combines prior information and makes full use of the characteristics of the sequential quadratic optimization algorithm with fast convergence speed and high accuracy, and can design ballistic parameters that meet the ballistic performance requirements of the launch vehicle and make the launch vehicle have greater launch environment adaptability. The method studied in this paper can provide technical support for the overall scheme demonstration of the launch vehicle, ballistic optimization design and in-atmosphere perturbation guidance research.

[0113] Based on the same inventive concept, the embodiment of the present application also provides a launch vehicle control device based on wind field characteristics corresponding to the launch vehicle control method based on wind field characteristics provided in the above embodiment. Since the principle of solving problems by the device in the embodiment of the present application is similar to that of the launch vehicle control method based on wind field characteristics in the above embodiment of the present application, the implementation of the device can refer to the implementation of the method, and the repeated parts will not be described again.

[0114] As Figure 7 shown, Figure 7It is a functional module diagram of a carrier rocket control device based on wind field characteristics provided in an embodiment of the present application. The carrier rocket control device 10 based on wind field characteristics includes: an acquisition module 101, which acquires a wind field model generated by fitting through actually measured wind field data; a determination module 102, which determines the relative velocity of the carrier rocket relative to the atmosphere at the current moment in the aircraft coordinate system according to the velocity of the carrier rocket at the current moment in the launch coordinate system, the wind speed at the current moment simulated in the wind field model in the launch coordinate system, and a first conversion relationship, where the first conversion relationship is used to convert the launch coordinate system to the aircraft coordinate system; a first calculation module 103, which calculates an influence parameter through the relative velocity expression and calculates the aerodynamic force at the current moment according to the influence parameter; a second calculation module 104, which integrates the differential equations of the centroid motion of the carrier rocket by combining the aerodynamic force, a preset initial value of a preset control parameter, and the flight program angle expression of the carrier rocket changing with time, and calculates the preset state variables of the carrier rocket at the next moment, and cycles in this way to obtain the preset state variables at each moment; a third calculation module 105, which calculates the objective function of the carrier rocket according to the preset state variables at each moment; an optimization module 106, which based on the sequential quadratic programming algorithm, changes the value of the preset control parameter through the objective function and preset constraint conditions, and calculates the corresponding objective function under different preset control parameters to determine the optimized value of the preset control parameter.

[0115] Based on the same inventive concept, refer to Figure 8 As shown, it is a schematic structural diagram of an electronic device provided in an embodiment of the present application. The electronic device 20 includes: a processor 201, a memory 202, and a bus 203. The memory 202 stores machine-readable instructions executable by the processor 201. When the electronic device 20 runs, communication is carried out between the processor 201 and the memory 202 through the bus 203. When the machine-readable instructions are run by the processor 201, the steps of the carrier rocket control method based on wind field characteristics as described in any one of the above embodiments are executed.

[0116] Specifically, when the machine-readable instructions are executed by the processor 201, the following processes may be performed: obtaining a wind field model generated by fitting through actually measured wind field data; determining the relative velocity of the launch vehicle relative to the atmosphere at the current moment in the aircraft coordinate system according to the velocity of the launch vehicle at the current moment in the launch coordinate system, the wind speed at the current moment simulated in the wind field model in the launch coordinate system, and a first conversion relationship, where the first conversion relationship is used to convert the launch coordinate system to the aircraft coordinate system; calculating an influence parameter through the relative velocity, and calculating the aerodynamic force at the current moment according to the influence parameter; integrating the differential equations of the centroid motion of the launch vehicle by combining the aerodynamic force, the preset initial value of the preset control parameter, and the expression of the flight program angle of the launch vehicle changing with time, calculating the preset state variables of the launch vehicle at the next moment, and circulating in this way to obtain the preset state variables at each moment; calculating the objective function of the launch vehicle according to the preset state variables at each moment; based on the sequential quadratic programming algorithm, changing the value of the preset control parameter through the objective function and the preset constraint conditions, and calculating the corresponding objective function under different preset control parameters to determine the optimized value of the preset control parameter.

[0117] Based on the same inventive concept, an embodiment of the present application further provides a computer-readable storage medium, on which a computer program is stored, and when the computer program is run by a processor, the steps of the method for controlling a launch vehicle based on wind field characteristics provided in the above embodiment are executed.

[0118] Specifically, the storage medium can be a general storage medium, such as a removable disk, a hard disk, etc. When the computer program on the storage medium runs, it can execute the above-mentioned control method for a launch vehicle based on wind field characteristics. By means of a wind field model constructed according to the actually measured wind field data, the relative velocity of the launch vehicle relative to the atmosphere in the aircraft coordinate system is determined according to the wind speed simulated by the wind field model, and the aerodynamic force is calculated through the relative velocity. Then, in combination with the aerodynamic force, the preset initial value of the preset control parameter, and the expression of the flight program angle of the launch vehicle changing with time, the differential equations of the centroid motion of the launch vehicle are integrated to calculate the preset state variables of the launch vehicle at each moment, and the objective function of the preset state variables is output. The sequential quadratic programming algorithm is used to change the values of different preset control parameters in combination with the objective function and the preset constraint conditions, and the objective functions corresponding to different preset control parameters are calculated respectively to determine the optimized values of the preset control parameters, so as to control the launch vehicle to execute the flight mission according to the optimized values of the preset control parameters, solving the technical problem in the prior art that the influence of the wind field on the launch vehicle is not considered, and achieving the technical effect of improving the success probability of the launch vehicle executing the flight mission. That is to say, in the ballistic design stage, usually considering factors such as the measurement level and the influence magnitude, the random factor influence is ignored. The present application actively considers and takes the wind vector as a constant to analyze the influence on the ballistic design.

[0119] Those skilled in the art can clearly understand that for the convenience and simplicity of description, the specific working processes of the above-described system and device can refer to the corresponding processes in the foregoing method embodiments and will not be elaborated herein. In the several embodiments provided in the present application, it should be understood that the disclosed system, device, and method can be implemented in other ways. The device embodiments described above are merely illustrative. For example, the division of the units is only a logical function division, and there may be other division methods in actual implementation. For another example, multiple units or components can be combined or integrated into another system, or some features can be ignored or not executed. Another point is that the displayed or discussed coupling or direct coupling or communication connection to each other can be through some communication interfaces. The indirect coupling or communication connection of the device or unit can be in an electrical, mechanical, or other form.

[0120] The units described as separate components may or may not be physically separated. The components displayed as units may or may not be physical units, that is, they may be located in one place, or may be distributed to multiple network units. Some or all of the units can be selected according to actual needs to achieve the purpose of the solution of this embodiment.

[0121] In addition, in each embodiment of the present application, each functional unit may be integrated into one processing unit, may exist physically alone for each unit, or two or more units may be integrated into one unit.

[0122] If the above-mentioned functions are implemented in the form of software functional units and sold or used as independent products, they can be stored in a non-volatile computer-readable storage medium executable by a processor. Based on such an understanding, the technical solution of the present application essentially, or the part that contributes to the prior art, or part of the technical solution, can be embodied in the form of a software product. The computer software product is stored in a storage medium and includes several instructions for causing a computer device (which may be a personal computer, a server, or a network device, etc.) to execute all or part of the steps of the methods described in each embodiment of the present application. The foregoing storage medium includes: various media such as USB flash drives, mobile hard disks, read-only memory (ROM), random access memory (RAM), magnetic disks, or optical discs that can store program codes.

[0123] The above is only the specific implementation manner of the present application, but the protection scope of the present application is not limited thereto. Any person skilled in the art within the technical scope disclosed in the present application can easily think of changes or substitutions, which should all be covered by the protection scope of the present application. Therefore, the protection scope of the present application shall be subject to the protection scope of the claims.

Claims

1. A control method for a launch vehicle based on wind field characteristics, characterized in that, The method includes: Obtaining a wind field model generated by fitting through actually measured wind field data; Determining the relative velocity of the launch vehicle relative to the atmosphere at the current moment in the aircraft coordinate system according to the velocity of the launch vehicle at the current moment in the launch coordinate system, the wind speed at the current moment simulated in the wind field model, and a first conversion relationship, where the first conversion relationship is used to convert the launch coordinate system to the aircraft coordinate system; Calculating an influence parameter through the relative velocity, and calculating the aerodynamic force at the current moment according to the influence parameter; Integrating the differential equations of the centroid motion of the launch vehicle by combining the aerodynamic force, the preset initial value of the preset control parameter, and the flight program angle expression of the launch vehicle changing with time, calculating the preset state variable of the launch vehicle at the next moment, and obtaining the preset state variables at each moment in this way; Calculating the objective function of the launch vehicle according to the preset state variables at each moment; Based on the sequential quadratic programming algorithm, changing the value of the preset control parameter through the objective function and the preset constraint conditions, and calculating the corresponding objective function under different preset control parameters to determine the optimized value of the preset control parameter; Among them, the aerodynamic force is calculated in the following way: Calculating the influence parameter and the dynamic pressure that affect the aerodynamic coefficient through the velocities of the relative velocity under each coordinate axis of the aircraft coordinate system, where the influence parameter includes the angle of attack and the sideslip angle; Calculating the aerodynamic force according to the angle of attack, the sideslip angle, the aerodynamic area of the launch vehicle, the dynamic pressure, and the aerodynamic coefficient selected through the influence parameter; Among them, the flight program angle expression is used to describe the program angle in each time period during the flight of the launch vehicle, and the flight program angle expression includes: Among them, is the flight program angle varying with time, is the vertical flight section of the launch vehicle, is the ballistic inclination angle, is the angle of attack varying with time, is the turning section of the launch vehicle, is the preset program angle, is the constant flight program angle section of the launch vehicle, is from time to start maintaining a constant flight program angle slope, is the constant slope turning section of the launch vehicle, is the constant pitch program angle maintained by the launch vehicle before the shutdown point, is the aiming section of the launch vehicle in the second-stage flight phase.

2. The method according to claim 1, characterized in that, The preset control parameter includes the firing azimuth angle, the maximum value of the absolute value of the angle of attack in the turning section, the time when the angle of attack reaches the extreme value in the turning section, the slope of the flight program angle, and the starting time of the aiming section in the second-stage flight phase of the launch vehicle; Among them, the angle of attack changing with time is described by the following formula: wherein, is the angle of attack varying with time, is the maximum value of the absolute value of the angle of attack within the turning section, is a preset constant value, is the time when the angle of attack reaches an extreme value within the turning section.

3. The method according to claim 1, wherein The differential equations of the centroid motion of the launch vehicle are described by the following formula: Wherein, is the acceleration, is the second transformation relationship, and the second transformation relationship is used to transform the aircraft coordinate system to the launch coordinate system, is the apparent acceleration in the aircraft coordinate system, is the earth rotation speed, is the coordinate vector of the earth center in the launch coordinate system, is the position vector is the derivative with respect to time, is the velocity of the launch vehicle at each moment in the launch coordinate system, and g is the gravitational acceleration.

4. The method according to claim 1, characterized in that, The objective function of the launch vehicle is calculated in the following way: Calculating the target parameter value of the launch vehicle according to the preset state variables at each moment, where the target parameter value includes at least one of the following: the payload under the given orbit injection condition, the elliptical orbit under the given payload, and the range; Calculating the objective function of the launch vehicle according to each target parameter value and its corresponding preset weight factor.

5. The method according to claim 1, wherein The optimized value of the preset control parameter is determined in the following way: Based on the sequential quadratic programming algorithm, determining the current search direction and the current search step corresponding to the objective function and the preset constraint conditions; Updating the preset control parameter according to the current search direction and the current search step to obtain the updated value of the preset control parameter; Determine whether a preset convergence condition is satisfied. If the preset convergence condition is not satisfied, return to recalculate the preset control parameters of the launch vehicle until the preset convergence condition is satisfied, and use the updated value of the preset control parameter corresponding to when the convergence condition is reached as the optimized value of the preset control parameter.

6. A control device for a launch vehicle based on wind field characteristics, characterized in that, The device includes: An acquisition module that acquires a wind field model generated by fitting through actually measured wind field data; A determination module that determines the relative velocity of the launch vehicle relative to the atmosphere at the current moment in the aircraft coordinate system based on the velocity of the launch vehicle at the current moment in the launch coordinate system, the wind speed at the current moment simulated in the wind field model in the launch coordinate system, and a first conversion relationship, where the first conversion relationship is used to convert the launch coordinate system to the aircraft coordinate system; A first calculation module that calculates an influence parameter through the relative velocity and calculates the aerodynamic force at the current moment based on the influence parameter; A second calculation module that integrates the differential equations of the centroid motion of the launch vehicle by combining the aerodynamic force, the preset initial value of the preset control parameter, and the flight program angle expression of the launch vehicle varying with time, calculates the preset state variable of the launch vehicle at the next moment, and obtains the preset state variables at each moment in this way; A third calculation module that calculates the objective function of the launch vehicle based on the preset state variables at each moment; An optimization module that, based on the sequential quadratic programming algorithm, changes the value of the preset control parameter through the objective function and the preset constraint condition, and calculates the objective function corresponding to different preset control parameters to determine the optimized value of the preset control parameter; Among them, the first calculation module calculates the aerodynamic force in the following manner: Calculate the influence parameter and the dynamic pressure that affect the aerodynamic coefficient through the velocities of the relative velocity in each coordinate axis of the aircraft coordinate system, where the influence parameter includes the angle of attack and the sideslip angle; Calculate the aerodynamic force based on the angle of attack, the sideslip angle, the aerodynamic area of the launch vehicle, the dynamic pressure, and the aerodynamic coefficient selected through the influence parameter; Among them, the flight program angle expression is used to describe the program angle of each time period during the flight of the launch vehicle, and the flight program angle expression includes: Among them, is the flight program angle varying with time, is the vertical flight section of the launch vehicle, is the ballistic inclination angle, is the angle of attack varying with time, is the turning section of the launch vehicle, is the preset program angle, is the constant flight section of the program angle of the launch vehicle, is from time to maintain a constant slope of the flight program angle starting from, is the constant slope turning section of the launch vehicle, is the constant pitch program angle maintained by the launch vehicle before the shutdown point, is the aiming section of the launch vehicle in the second-stage flight phase.

7. An electronic device, characterized in that, Including: A processor, a memory, and a bus. The memory stores machine-readable instructions executable by the processor. When the electronic device runs, the processor communicates with the memory through the bus. When the machine-readable instructions are run by the processor, the steps of the launch vehicle control method based on wind field characteristics according to any one of claims 1 to 5 are executed.

8. A computer-readable storage medium, characterized in that, A computer program is stored on the computer-readable storage medium. When the computer program is run by the processor, the steps of the launch vehicle control method based on wind field characteristics according to any one of claims 1 to 5 are executed.

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