Carrier rocket control method based on wind field characteristics

Abstract

The invention provides a carrier rocket control method based on wind field characteristics, and the method comprises the steps: obtaining a wind field model generated through the fitting of actually measured wind field data; according to the speed of the carrier rocket at the current moment under the launching coordinate system, the wind speed at the current moment and the first conversion relation, the relative speed of the carrier rocket relative to the atmosphere at the current moment in the aircraft coordinate system is determined; influence parameters are calculated through the relative speed, and the aerodynamic force at the current moment is calculated according to the influence parameters; integrating the mass center motion differential equation set of the carrier rocket by combining the aerodynamic force, the preset initial value of the preset control parameter and the flight program angle expression of the carrier rocket changing along with time, and circularly calculating the preset state variable of each moment; calculating a target function of the carrier rocket according to the preset state variable at each moment; and on the basis of a sequential quadratic programming algorithm, target functions corresponding to different preset control parameters are calculated through the target functions and preset constraint conditions, so that optimal values of the preset control parameters are determined.

Description

Technical Field

[0001] The present application relates to the field of launch vehicle technology, and in particular to a launch vehicle control method 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 incoming 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 amount of deviation between the aerodynamic force and aerodynamic moment of the launch vehicle and the aerodynamic force and aerodynamic moment in the static atmosphere, which interferes 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 inability of the stability control system to respond accurately, making it difficult to control the missile to fly according to the predetermined trajectory plan. Especially for rocket launches under a static base, the missile speed in the initial stage is relatively low, and the influence of the wind vector cannot be ignored, which affects the normal implementation of the launch mission to a certain extent. Summary of the invention

[0003] In view of this, the purpose of the present application is to at least provide a launch vehicle control method based on wind field characteristics, by constructing a wind field model according to the actually measured wind field data, determining the relative speed 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 speed, and integrating the differential equations of the center of mass 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 that changes with time, so as to calculate the preset state variables of the launch vehicle at each moment, and output the objective function of the preset state variables, and change the values ​​of different preset control parameters by combining the objective function and the preset constraints through the sequential quadratic programming algorithm, and correspondingly calculate the objective functions corresponding to different preset control parameters, so as to determine the optimized values ​​of the preset control parameters, so as to facilitate the control of the launch vehicle to perform the flight mission according to the optimized values ​​of the preset control parameters, solve the technical problem that the influence of the wind field on the launch vehicle is not considered in the prior art, and achieve the technical effect of improving the success probability of the launch vehicle to perform the flight mission.

[0004] This 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. 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; integrating the differential equations of the centroid motion of the launch vehicle by 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 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 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 in 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 angle of the launch vehicle in each time period during the flight process. The flight program angle expression includes:

[0008]

[0009] where, is the flight program angle changing 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 changing 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 maintained constant starting from time t2, t3 ≤ t < t4 is the constant slope turning section of the launch vehicle, is the pitch program angle maintained constant 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.

[0010] Optionally, the preset control parameters include the shooting azimuth, the maximum 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 flight program angle slope and the start time of the aiming section of the launch vehicle in the second stage flight phase, wherein the angle of attack that changes with time is described by the following formula:

[0011] α(t)=-4×α M ×Z×(1-Z)

[0012]

[0013] Among them, α(t) is the angle of attack that changes with time, α M is the maximum absolute value of the angle of attack in the turning section, a is a preset constant value, t m It is the time when the angle of attack reaches the extreme value in the turning section.

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

[0015]

[0016] in, is the acceleration, is a second transformation relationship, wherein the second transformation relationship is used to transform the aircraft coordinate system into the launch coordinate system, is the apparent acceleration in the aircraft coordinate system, 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 time derivative of the position vector X, V is the velocity of the launch vehicle at each moment in the launch coordinate system, and g is the gravitational acceleration.

[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 target parameter value of the launch vehicle is calculated, and the target parameter value includes at least one of the following: the payload under given orbital conditions, the elliptical orbit under given load, and the range; based on each target 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 in the following manner: 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 condition 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, and if the preset convergence condition is not met, it is returned to recalculate the preset control parameter of the launch vehicle until the preset convergence condition is met, and the updated value of the preset control parameter corresponding to the convergence condition is used 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 comprising: an acquisition module, which acquires a wind field model generated by fitting actual measured wind field data; a determination module, which determines a relative velocity expression of the launch vehicle relative to the atmosphere in the aircraft coordinate system at the current moment according to the velocity of the launch vehicle in the launch coordinate system at the current moment, the wind speed in the launch coordinate system at the current moment simulated in the wind field model, and a first transformation relationship, wherein the first transformation relationship is used to convert the launch coordinate system to the aircraft coordinate system; a first calculation module, which calculates an influencing parameter through the relative velocity expression, and calculates the air velocity at the current moment according to the influencing parameter. aerodynamic force; a second calculation module, combining the aerodynamic force, the preset initial value of the preset control parameter and the flight program angle expression of the launch vehicle that changes with time, integrates the differential equation group of the center of mass motion of the launch vehicle, calculates the preset state variables of the launch vehicle at the next moment, and obtains the preset state variables at each moment in a cycle; a third calculation module, calculates the objective function of the launch vehicle according to the preset state variables at each moment; an optimization module, based on a sequential quadratic programming algorithm, changes the value of the preset control parameter through the objective function and preset constraints, and calculates the corresponding objective function under 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, comprising: a processor, a memory and a bus, wherein the memory stores machine-readable instructions executable by the processor, and when the electronic device is running, the processor and the memory communicate through the bus, and the machine-readable instructions are executed by the processor to execute the steps of the launch vehicle control method based on wind field characteristics described in the first aspect or any possible implementation manner of the first aspect.

[0021] In a fourth aspect, 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 executed by a processor, the steps of the launch vehicle control method based on wind field characteristics described in the above-mentioned first aspect or any possible implementation manner of the first aspect are executed.

[0022] A launch vehicle control method based on wind field characteristics is provided in an embodiment of the present application. The method comprises: obtaining a wind field model generated by fitting actual measured wind field data; determining the relative speed of the launch vehicle relative to the atmosphere in the aircraft coordinate system at the current moment according to the speed of the launch vehicle in the launch coordinate system at the current moment, the wind speed in the launch coordinate system at the current moment simulated in the wind field model, and a first conversion relationship, wherein the first conversion relationship is used to convert the launch coordinate system to the aircraft coordinate system; calculating an influencing parameter through the relative speed, and calculating the aerodynamic force at the current moment according to the influencing parameter; In combination with the aerodynamic force, the preset initial values ​​of the preset control parameters and the flight program angle expression of the launch vehicle that changes with time, the differential equation group of the center of mass motion of the launch vehicle is integrated to calculate the preset state variables of the launch vehicle at the next moment, and the preset state variables at each moment are obtained in a cycle; the objective function of the launch vehicle is calculated based on the preset state variables at each moment; based on the sequential quadratic programming algorithm, the values ​​of the preset control parameters are changed through the objective function and preset constraints, and the corresponding objective functions under different preset control parameters are calculated to determine the optimized values ​​of the preset control parameters. By constructing a wind field model according to the actually measured wind field data, the relative speed of the carrier rocket 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 by the relative speed. The differential equation group of the center of mass motion of the carrier rocket is integrated by combining the aerodynamic force, the preset initial value of the preset control parameter and the flight program angle expression of the carrier rocket that changes with time, so as to calculate the preset state variables of the carrier rocket at each moment, and output the objective function of the preset state variables. The values ​​of different preset control parameters are changed by combining the objective function and the preset constraint conditions through the sequential quadratic programming algorithm, and the objective functions corresponding to different preset control parameters are calculated respectively, so as to determine the optimized values ​​of the preset control parameters, so as to facilitate the control of the carrier rocket to perform the flight mission according to the optimized values ​​of the preset control parameters, solve the technical problem that the influence of the wind field on the carrier rocket is not considered in the prior art, and achieve the technical effect of improving the success probability of the carrier rocket to perform the flight mission.

[0023] In order to make the above-mentioned objects, features and advantages of the present application more obvious and easy to understand, preferred embodiments are specifically cited below and described in detail with reference to the attached drawings. BRIEF DESCRIPTION OF THE DRAWINGS

[0024] In order to more clearly illustrate the technical solutions of the embodiments of the present application, the drawings required for use in the embodiments will be briefly introduced below. It should be understood that the following drawings only show certain embodiments of the present application and therefore should not be regarded as limiting the scope. For ordinary technicians in this field, other related drawings can be obtained based on these drawings without paying creative work.

[0025] Figure 1 A flow chart of a launch vehicle control method based on wind field characteristics provided in an embodiment of the present application is shown.

[0026] Figure 2 A curve chart showing the change of wind speed with height in the wind field model provided in an embodiment of the present application is shown.

[0027] Figure 3 A curve showing the change in wind speed with height of the wind field model provided in an embodiment of the present application is shown.

[0028] Figure 4 A curve showing the change in wind direction with height of the wind field model provided in an embodiment of the present application is shown.

[0029] Figure 5 A design curve diagram of the program angle corresponding to the wind speed considering the wind field model and the wind speed not considering the wind field model provided in the embodiment of the present application is shown.

[0030] Figure 6 A schematic diagram showing the maximum aerodynamic force corresponding to a wind field model considered and the maximum aerodynamic force corresponding to a wind field model not considered provided in an embodiment of the present application is shown.

[0031] Figure 7 A functional module diagram of a launch vehicle control device based on wind field characteristics provided in an embodiment of the present application is shown.

[0032] Figure 8 A schematic structural diagram of an electronic device provided in an embodiment of the present application is shown. DETAILED DESCRIPTION

[0033] To make the purpose, technical scheme and advantages of the embodiments of the present application clearer, the technical scheme in the embodiments of the present application will be clearly and completely described below in conjunction with the drawings in the embodiments of the present application. It should be understood that the drawings in the present application only serve the purpose of explanation and description and are not used to limit the scope of protection of the present application. In addition, it should be understood that the schematic drawings are not drawn in real proportion. The flowchart used in this application shows the operations implemented according to some embodiments of the present application. It should be understood that the operations of the flowchart can be implemented out of sequence, and the steps without logical context can be reversed in order or implemented simultaneously. In addition, those skilled in the art, under the guidance of the content of the present application, can add one or more other operations to the flowchart, or remove one or more operations from the flowchart.

[0034] In addition, the described embodiments are only a part of the embodiments of the present application, rather than all the embodiments. The components of the embodiments of the present application described and shown in the drawings here can be arranged and designed in various configurations. Therefore, the following detailed description of the embodiments of the present application provided in the drawings is not intended to limit the scope of the application claimed for protection, but merely represents the selected embodiments of the present application. Based on the embodiments of the present application, all other embodiments obtained by those skilled in the art without making creative work belong to the scope of protection of the present application.

[0035] In the prior art, trajectory design plays an extremely important role in the overall design of a launch vehicle, and it runs through the entire process of the overall design. The overall plan, design parameters, combat performance and flight plan of the rocket are closely related to the flight trajectory. From the perspective of trajectory, theoretical research and launch tests have shown that the deviation of atmospheric parameters from their standard values ​​has a great influence on the movement of the launch vehicle. In traditional trajectory design, it is customary to take the speed of the launch vehicle in the inertial coordinate system as the incoming flow speed 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 the aerodynamic force and aerodynamic moment in the static atmosphere, which interferes 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 inability of the stability control system to respond accurately, and it is difficult to control the missile to fly according to the predetermined trajectory plan. Especially for rocket launches under a static base, the missile speed is low in the initial stage, and the influence of the wind vector cannot be ignored, which affects the normal implementation of the launch mission to a certain extent.

[0036] Based on this, the embodiment of the present application provides a launch vehicle control method based on wind field characteristics, by constructing a wind field model according to the actual measured wind field data, determining the relative speed 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 speed, and integrating the differential equations of the center of mass 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 over time, so as to calculate the preset state variables of the launch vehicle at each moment, and output the objective function of the preset state variables, and change the values ​​of different preset control parameters by combining the objective function and the preset constraint conditions through the sequential quadratic programming algorithm, and correspondingly calculating the objective functions corresponding to different preset control parameters, so as to determine the optimized values ​​of the preset control parameters, so as to control the launch vehicle to perform the flight mission according to the optimized values ​​of the preset control parameters, and solve the technical problem that the influence of the wind field on the launch vehicle is not considered in the prior art, and achieve the technical effect of improving the success probability of the launch vehicle to perform the flight mission. The details are as follows:

[0037] See also Figure 1 , Figure 1 The following is a flow chart of a launch vehicle control method based on wind field characteristics provided in an embodiment of the present application. Figure 1 As shown, the launch vehicle control method based on wind field characteristics provided in the embodiment 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, wind field data includes wind speed and wind direction. Although the atmospheric wind field has different shapes and forms, and will change with factors such as terrain, longitude and latitude, altitude, temperature, air density, time, etc. However, long-term research on the wind field at a fixed location has found that the change of the wind field shows obvious statistical significance characteristics. That is, according to the effect of wind on the movement of the launch vehicle, the wind speed can be regarded as three components with different characteristics, including steady wind, shear wind and gust. Steady wind refers to horizontal wind that changes very slowly relative to time. The curve of this type of wind changing with altitude in a certain place and season can be obtained by statistics of a large amount of observation data. Shear wind is horizontal wind that only appears at a certain altitude. Its wind speed increases sharply with increasing altitude, and then decreases sharply. 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 short-term wind force, just like pulse torque. Gust refers to the local disturbed airflow in the air around the launch vehicle, which can usually be regarded as a steady random process and described by spectral density.

[0040] For example, see Figure 2 , Figure 2 This is a graph showing the wind speed changing with height in the wind field model provided in the embodiment of the present application. Figure 2 As shown, the horizontal coordinate of the curve of wind speed varying with height in the wind field model provided in the embodiment of the present application is wind speed, in meters per second, and the vertical coordinate is height, in 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, the wind speed curve a can be understood as being obtained by merging the steady wind speed curve b, the maximum wind speed envelope curve c and the shear wind curve d, and the wind speed curve a can also include gusts. Among them, different wind speed curves have different rates of change with height, the steady wind speed curve has the longest period of change with height, the shear wind is second, and the gusts are the shortest.

[0041] That is to say, the wind field at each flight time can be obtained by linearly superimposing the gusts and the maximum steady wind speed, and the shear wind is only considered in the most serious cases. Furthermore, the fixed location is selected as the launch site for the carrier rocket, and the wind speed and wind direction of the launch site for the carrier rocket are detected for a long time, and a wind field model that produces different wind speed and wind direction changes over time is fitted by the detected wind speed and wind direction, and the wind field model is used to simulate the wind speed and wind direction at each time. Moreover, the wind field model shows statistical periodic changes with the seasons. Exemplarily, by collecting the wind speed and wind direction in fixed months every year 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 relative velocity expression of the carrier rocket relative to the atmosphere in the aircraft coordinate system at the current moment based on the velocity of the carrier rocket in the launch coordinate system at the current moment, the wind speed in the launch coordinate system at the current moment simulated in the wind field model, and the first conversion relationship.

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

[0044] Among them, the aircraft refers to the launch vehicle in this application. Furthermore, when the flight mission of the launch vehicle is known, the launch coordinate system can be constructed by knowing the launch aiming direction of the launch vehicle, and the aircraft coordinate system can be constructed by the center of mass, longitudinal axis and longitudinal symmetry plane of the launch vehicle, and the first transformation relationship of converting the launch coordinate system to the aircraft coordinate system can be known.

[0045] Exemplarily, the wind vector at a certain moment in the wind field model is expressed in the emission 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 Celestial East coordinate system to the launch coordinate system, W E is the wind vector at each moment of the launch site of the carrier rocket in the wind field model in the north sky east coordinate system, a0 is the wind direction, a c is the wind speed constant. The origin of the North Sky East coordinate system can be the same as the origin of the launch coordinate system, or a random origin can be selected, with the X axis pointing to the north, the Y axis pointing to the sky, and the Z axis pointing to the east.

[0048] That is to say, the wind field model can directly simulate the wind vector at each time in the transmission coordinate system, or it can simulate the wind vector at each time in the North Sky East coordinate system and then convert the North Sky East coordinate system to the transmission coordinate system through the conversion matrix to calculate the wind vector at a certain time in the transmission coordinate system. This application does not limit the coordinate system used in the wind field model, and only needs to determine the wind vector in the transmission coordinate system through the conversion matrix.

[0049] Furthermore, taking the wind vector into account, the velocity of the launch vehicle relative to the atmosphere in the launch coordinate system is expressed by the following formula:

[0050] V W =VW (2)

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

[0052] That is to say, by calculating the speed of the carrier rocket at the current moment in the launch coordinate system in step S104, the speed of the carrier rocket relative to the atmosphere at the current moment in the launch coordinate system can be calculated in combination with the wind vector at the current moment in the wind field model. Therefore, the subsequent steps are executed by integrating the differential equations of the center of mass motion of the carrier rocket at the current moment, and the speed of the carrier rocket at the next moment in the launch coordinate system is solved, and step S102 is re-executed to realize the cycle, and the preset state variables of the carrier rocket at each time are calculated. In addition, the initial speed of the carrier rocket in the launch coordinate system is known.

[0053] Therefore, the relative velocity expression of the launch vehicle relative to the atmosphere in the vehicle coordinate system can be determined by 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:

[0054]

[0055] In formulas (3) and (4), V1 is the relative velocity of the launch vehicle relative to the atmosphere in the spacecraft coordinate system at each moment, V x1 is the relative velocity of the launch vehicle on the X axis of the vehicle coordinate system relative to the atmosphere at each moment, V y1 is the relative velocity of the launch vehicle on the Y axis of the vehicle coordinate system relative to the atmosphere at each moment, V z1 is the relative velocity of the launch vehicle on the Z axis of the vehicle coordinate system relative to the atmosphere at each moment, V is the transformation matrix from the launch coordinate system to the aircraft coordinate system. 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, and γ is the roll program angle. The other two program corners can be ignored and will be discussed later in this application. To represent the pitch program angle.

[0056] That is to say, the present application realizes passive load reduction by taking into account the influence of the wind field on the launch vehicle and thereby calculating the speed and aerodynamic force of the launch vehicle relative to the atmosphere.

[0057] S103: Calculate an influencing parameter using the relative speed, and calculate the aerodynamic force at the current moment based on the influencing parameter.

[0058] How to determine the aerodynamic force acting on a launch vehicle when it moves relative to the atmosphere is a rather complicated problem, and it is difficult to accurately determine it through theoretical calculations. Currently, a method is used that combines calculations based on aerodynamic theory with corrections based on aerodynamic experiments. Aerodynamic experiments are conducted in a wind tunnel that can produce uniform airflow at a certain Mach number. For example, a physical model of a scaled-down launch vehicle is obtained, and the physical model is placed in a wind tunnel for a wind tunnel test, so that airflow blows through the model at a certain Mach number. At this time, the model is placed in the wind tunnel and has no power. It is only based on the speed V of the launch vehicle relative to the atmosphere in the launch coordinate system. W To simulate, and try to meet the requirements of each V WThe corresponding aerodynamic coefficients under multiple sideslip angle ranges and multiple angle of attack ranges are calculated by measuring the aerodynamic forces acting on the model, and then the aerodynamic forces acting on the real object under these Mach numbers are obtained by applying the similarity transformation principle. Among them, aerodynamic forces are divided into lift, drag and side force.

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

[0060] Specifically, the aerodynamic force is calculated in the following manner: the influencing parameters and dynamic pressure are calculated through the speed at each coordinate axis of the aircraft coordinate system in the relative speed, and the influencing parameters include the angle of attack and the sideslip angle; the aerodynamic force is 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 influencing parameters.

[0061] Among them, the influencing parameters are calculated by the following formula:

[0062]

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

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

[0065] The dynamic pressure is calculated by the following formula:

[0066]

[0067] In formula (7), q is the dynamic pressure, ρ is the air density, V h is the combined velocity of the launch vehicle in the spacecraft coordinate system, which can be obtained by x1 、V y1 、V z1 Sum the corresponding squares and then take the square root to get it.

[0068] The aerodynamic force is calculated by the following formula:

[0069]

[0070] In formula (8), A is the drag, N is the lift, Z1 is the lateral force, and C A is the aerodynamic coefficient corresponding to the drag, 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, 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 the range determined by wind tunnel test.

[0071] That is to say, the influencing parameters are used to influence the selection of aerodynamic coefficients, and the influencing parameters can also be reported as parameters such as altitude and Mach. At the current time, by determining the speed of the launch vehicle at each coordinate axis of 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 numerical table, curve or mathematical model under the constraints of different variables such as altitude, Mach, angle of attack, sideslip angle, etc. obtained in advance, and calculate the dynamic pressure through the speed of the launch vehicle at each coordinate axis of the aircraft coordinate system, so as to calculate 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 by the relative speed of the launch vehicle relative to the atmosphere in the aircraft coordinate system at the current moment, and then the aerodynamic coefficients corresponding to the angle of attack, sideslip angle and their ranges are used to calculate the aerodynamic force at the current moment through the angle of attack, sideslip angle, dynamic pressure and aerodynamic coefficients. Furthermore, the aerodynamic force can be calculated in this way by knowing the relative speed of the launch vehicle relative to the atmosphere in the aircraft coordinate system at each moment.

[0073] S104: 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 that changes with time, the differential equation group of the center of mass motion of the launch vehicle is integrated to calculate the preset state variables of the launch vehicle at the next moment, and the preset state variables at each moment are obtained in a cycle.

[0074] The flight program angle expression is used to describe the program angle of the launch vehicle in each time period during the flight. The preset initial value of the preset control parameter can be set by a technician in combination with experimental data, experience, and formulas.

[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 of the passive section are also determined. Generally, 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 nearly shut down.

[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, it is required that the rocket maintain a constant program angle during engine startup, shutdown, and before and after stage separation, and try 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 trajectory 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 require a launch pad with a simple structure, and at the same time can also ensure the stability of the launch vehicle at the takeoff moment. 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 programmed angle. The selection of t1 should be reasonable. If the time of the vertical flight section is too long, it will increase the gravitational loss of speed and require a large normal force during turning due to excessive speed; 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, generally, 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 time 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 time is calculated through the following formula:

[0081]

[0082] Among them, in formula (10), t1 is the end time 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 as 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 time 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. It can be obtained from formula (11) that α(t) starts to rapidly reach a negative value, and then the absolute value starts to decrease, approaching zero at an exponential rate. The rate of approaching zero is determined by the parameter a. By taking the derivative of α(t), we get When the slope of α(t) approaches 0, that is we get Substituting it into formula (12), we can obtain the preset constant value 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 for the purpose of facilitating the separation of the first and second stages of the launch vehicle and reducing the influence of interference during separation on the movement of the missile, and at the same time enabling the engine in the second-stage flight phase to reach the rated working 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 for 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 vehicle 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, aerodynamics, 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 time derivative of the position vector X, V is the velocity of the launch vehicle at each moment in the launch coordinate system, and g is the gravitational acceleration.

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

[0095] The fourth-order Runge-Kutta method calculates the value of the next point by using the weighted average of the four slopes within the time step, thereby calculating the coordinates and speed of the launch vehicle in the spacecraft 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 preset initial values ​​of the preset control parameters, the coordinates and speed at the next moment are solved through the aerodynamic force and program angle at the current moment, and the coordinates and speed at the next moment are used as the new current moment for looping, so as to solve the coordinates and speed corresponding to each moment.

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

[0098] Specifically, the objective function of the launch vehicle is calculated in the following manner: based on the preset state variables at each moment, the target parameter value of the launch vehicle is calculated, and the target parameter value includes at least one of the following: the payload under given orbital conditions, the elliptical orbit under given load, and the range; based on each target parameter value and its corresponding preset weight factor, the objective function of the launch vehicle is calculated.

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

[0100] For example, given orbital conditions are generally the orbital roots, orbital altitude and orbital velocity at the time of orbital entry, which correspond to preset values. It is only necessary to change the mass of the launch vehicle under different payloads. The mass of the launch vehicle m gis the sum of the payload mass and the structure mass, which is used to calculate the preset state variables at different launch vehicle masses at different times, so as to determine the launch vehicle mass that meets the given orbital conditions, and thus obtain the payload under the given orbital conditions corresponding to the preset control parameters as the preset initial values. Among them, the launch vehicle mass is related to the apparent acceleration in the calculation vehicle coordinate system.

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

[0102] Furthermore, according to the calculation method of each target parameter value, each target parameter value is calculated by the coordinates and speed at each time corresponding to the preset initial value of the preset control parameter, and then according to the preset weight factors of each target parameter value, the value of the objective function of the launch vehicle is calculated by summing the product of each target parameter value and the preset weight factor. This objective function value represents the objective function value corresponding to the preset initial value of the preset control parameter, and the 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 parameters are changed through the objective function and the preset constraints, and the corresponding objective functions under different preset control parameters are calculated to determine the optimized values ​​of the preset control parameters.

[0104] Specifically, the optimal value of the preset control parameter is determined in the following manner: 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 condition 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, and if the preset convergence condition is not met, it is returned to recalculate the preset control parameter of the launch vehicle until the preset convergence condition is met, and the updated value of the preset control parameter corresponding to the convergence condition is used as the optimal value of the preset control parameter.

[0105] Among them, the preset constraints include landing point constraints, preset control parameter constraints, dynamic pressure constraints and normal overload constraints. Among them, the landing point constraint refers to the constraint of the terminal ballistic parameters of the launch vehicle, such as the orbital root number constraint of the launch vehicle's entry point. The preset control parameter constraint is used to describe the value range of each preset control parameter. For example, 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 optimization value of each preset control parameter should belong to its corresponding value range. From the perspective of thermal protection system design, the dynamic pressure should be within a certain range to ensure that the surface insulation material structure is not damaged, the control hinge torque increases with the increase of dynamic pressure, and the dynamic pressure should also be kept within the dynamic pressure allowed by the maximum hinge torque required by the control surface. At the same time, limiting the dynamic pressure can also ensure that the launch vehicle maintains stability in lateral flight to a certain extent. For example, the dynamic pressure should be less than or equal to the preset dynamic pressure value. 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 onboard equipment. In order 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., normal), m is the mass of the launch vehicle, and g is the gravitational acceleration.

[0108] Among them, the preset constraints can be changed according to the different design tasks of the launch vehicle. For example, the debris landing area constraints at all levels, the height constraints of the shutdown point during the flight test, and the aerodynamic thermal constraints are required. Furthermore, the preset constraints can be expressed as equality constraints and inequality constraints. The landing point constraint is an equality constraint, and the preset control parameter constraint, dynamic pressure constraint, and normal overload constraint are inequality constraints.

[0109] That is to say, a quadratic programming subproblem of the sequential quadratic programming algorithm is constructed according to the objective function and preset constraints corresponding to the preset initial value of the preset control parameter, and the direction and step size are searched by solving the quadratic programming subproblem, and then the value of the preset control parameter is updated by searching the direction and compensation to obtain an updated value of the preset control parameter, and the updated value can meet the preset constraints, and the coordinates and speeds at each time corresponding to the updated value are cyclically calculated in the manner of steps S102 to S104 through the updated value of the preset control parameter, and the objective function value is calculated in step S105; and, it is determined whether the preset convergence condition is met, and the preset convergence condition refers to The difference between the updated values ​​of the preset control parameters calculated by two iterative calculations conforms to the preset difference range. If the preset convergence condition is met, the optimized value of the preset control parameter is directly obtained. Alternatively, after the preset convergence condition is met, the updated value corresponding to the maximum effective payload under given orbital conditions, the maximum elliptical orbit under a given load, and the maximum range can be determined from each updated value as the optimized value of the preset control parameter; if the preset convergence condition is not met, the updated value of the preset control parameter is modified to the preset initial value, the search direction and the search step are recalculated to determine the updated value of the preset control parameter, and the fourth-order Runge-Kutta method is used to re-integrate the center of mass motion differential equation group to obtain the preset state variable, so as to perform a cyclic calculation.

[0110] For example, in order to specifically analyze the effect of passive load reduction trajectory design, firstly, a wind field model is constructed according to the historical wind field data of the aircraft launch site. Figure 3 and Figure 4 , Figure 3 This is a curve showing the wind speed versus height of the wind field model provided in the embodiment of the present application, where the horizontal axis is the wind speed in meters per second and the vertical axis is the height in kilometers. Figure 4 The curve of wind direction versus height of the wind field model provided in the embodiment of the present application, the horizontal axis is the wind direction angle, the unit is degree, and the vertical axis is the height, the unit is kilometer. Figure 3 and Figure 4 As shown, the wind direction angle is concentrated between 260° and 280°, the wind direction is westward, and the steady wind speed does not exceed 40m / s.

[0111] Since the partial influence of the wind field model is taken into account, for example, 30% of the wind speed influence is taken into account, 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 used in the actual calculation. Figure 5 , Figure 5 The embodiment of the present application provides a design curve diagram of the program angle corresponding to the wind speed with and without considering the wind field model. The horizontal axis is used to describe the flight time of the launch vehicle, and the vertical axis is used to describe the program angle. Figure 5As shown in the figure, the design of the program angle does not produce a significant change. On this basis, the maximum aerodynamic force of the launch vehicle flight under two conditions is simulated by considering the wind speed of the wind field model and ignoring the wind speed of the wind field model. Figure 6 , Figure 6 This is a schematic diagram of 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 in the embodiment of the present application. Figure 6 As shown in the figure, the maximum aerodynamic force is significantly reduced when the wind field model is considered to achieve passive load reduction. This reflects that under the same wind field, the load reduction trajectory can make the rocket body structure withstand greater high-altitude wind interference, greatly improving the launch environment adaptability of the carrier rocket and increasing the launch probability.

[0112] This application constructs a wind field model based on the wind field parameters of the existing launch site of the carrier rocket to reflect the wind field data under statistical characteristics. With the purpose of improving the launch probability, the motion equation under the influence of the wind field model is given in combination with the actual mission needs of the carrier rocket. Combined with multiple preset constraints and the program angle of planning, the numerical solution theory of trajectory optimization is studied, that is, the initial value of the control parameter is given based on practical experience, and the nonlinear equation with constraints is solved by the sequential quadratic programming method 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 fast convergence speed and high precision of the sequential quadratic optimization algorithm to design ballistic parameters that meet the ballistic performance requirements of the carrier rocket and make the carrier rocket have greater adaptability to the launch environment. The method studied in this paper can provide technical support for the overall program demonstration of the carrier rocket, trajectory optimization design and perturbation guidance research in the atmosphere.

[0113] Based on the same application concept, the embodiments of the present application also provide 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 embodiments. Since the principle of solving the problem by the device in the embodiments of the present application is similar to the launch vehicle control method based on wind field characteristics in the above embodiments of the present application, the implementation of the device can refer to the implementation of the method, and the repeated parts will not be repeated.

[0114] like Figure 7 As shown, Figure 7A functional module diagram of a launch vehicle control device based on wind field characteristics provided in an embodiment of the present application. The launch vehicle control device 10 based on wind field characteristics includes: an acquisition module 101, which acquires a wind field model generated by fitting actual measured wind field data; a determination module 102, which determines the relative speed of the launch vehicle relative to the atmosphere in the aircraft coordinate system at the current moment based on the speed of the launch vehicle in the launch coordinate system at the current moment, the wind speed in the launch coordinate system at the current moment simulated in the wind field model, and a first conversion relationship, wherein the first conversion relationship is used to convert the launch coordinate system to the aircraft coordinate system; a first calculation module 103, which calculates the influencing parameters through the relative speed expression, and calculates the aerodynamic force at the current moment based on the influencing parameters; a second calculation module 104, integrating the differential equations of the center of mass motion of the launch vehicle in combination with the aerodynamic force, the preset initial values ​​of the preset control parameters and the expression of the flight program angle of the launch vehicle that changes with time, calculating the preset state variables of the launch vehicle at the next moment, and thereby looping to obtain the preset state variables at each moment; a third calculation module 105, calculating the objective function of the launch vehicle according to the preset state variables at each moment; an optimization module 106, based on a sequential quadratic programming algorithm, changing the values ​​of the preset control parameters through the objective function and preset constraints, and calculating the corresponding objective functions under different preset control parameters to determine the optimized values ​​of the preset control parameters.

[0115] Based on the same application concept, see Figure 8 As shown, it is a schematic diagram of the structure 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 is running, the processor 201 and the memory 202 communicate through the bus 203, and the machine-readable instructions are executed by the processor 201 when it is running to perform the steps of the launch vehicle control method based on wind field characteristics as described in any of the above embodiments.

[0116] Specifically, when the machine-readable instructions are executed by the processor 201, the following processing can be performed: obtaining a wind field model fitted by actually measuring wind field data; determining the relative speed of the carrier rocket relative to the atmosphere in the aircraft coordinate system at the current moment according to the speed of the carrier rocket in the launch coordinate system at the current moment, the wind speed in the launch coordinate system at the current moment simulated in the wind field model, and a first conversion relationship, wherein the first conversion relationship is used to convert the launch coordinate system to the aircraft coordinate system; calculating the influencing parameters through the relative speed, and calculating the aerodynamic force at the current moment according to the influencing parameters; and The method comprises the following steps: integrating the differential equations of the center of mass motion of the launch vehicle by combining the aerodynamic force, the preset initial values ​​of the preset control parameters and the expression of the flight program angle of the launch vehicle that changes with time, calculating the preset state variables of the launch vehicle at the next moment, and cyclically obtaining the preset state variables at each moment; calculating the objective function of the launch vehicle according to the preset state variables at each moment; and changing the values ​​of the preset control parameters through the objective function and preset constraints based on a sequential quadratic programming algorithm, and calculating the corresponding objective functions under different preset control parameters to determine the optimized values ​​of the preset control parameters.

[0117] Based on the same application concept, an embodiment of the present application also provides a computer-readable storage medium, on which a computer program is stored. When the computer program is executed by a processor, the steps of the launch vehicle control method 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 mobile disk, a hard disk, etc. When the computer program on the storage medium is run, the above-mentioned launch vehicle control method based on wind field characteristics can be executed. The relative speed of the launch vehicle relative to the atmosphere in the aircraft coordinate system is determined according to the wind field model constructed according to the actually measured wind field data and the wind speed simulated by the wind field model, and the aerodynamic force is calculated by the relative speed. The differential equation group of the center of mass motion of the launch vehicle is integrated by combining the aerodynamic force, the preset initial value of the preset control parameter and the flight program angle expression of the launch vehicle that changes with time, so as to calculate the preset state variables of the launch vehicle at each moment, and output the objective function of the preset state variables. The values ​​of different preset control parameters are changed by combining the objective function and the preset constraints through the sequential quadratic programming algorithm, and the objective functions corresponding to different preset control parameters are calculated respectively, so as to determine the optimized values ​​of the preset control parameters, so as to facilitate the control of the launch vehicle to perform the flight mission according to the optimized values ​​of the preset control parameters, solve the technical problem that the influence of the wind field on the launch vehicle is not considered in the prior art, and achieve the technical effect of improving the success probability of the launch vehicle to perform the flight mission. That is to say, in the trajectory design stage, factors such as measurement level and impact size are usually considered, and the influence of random factors is ignored. This application actively considers and takes the wind vector as a constant to analyze the impact on the trajectory design.

[0119] Those skilled in the art can clearly understand that, for the convenience and simplicity of description, the specific working process of the system and device described above can refer to the corresponding process in the aforementioned method embodiment, and will not be repeated here. 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 schematic. For example, the division of the units is only a logical function division. There may be other division methods in actual implementation. For 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 mutual coupling or direct coupling or communication connection shown or discussed can be through some communication interfaces, indirect coupling or communication connection of devices or units, which can be electrical, mechanical or other forms.

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

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

[0122] If the function is implemented in the form of a software functional unit and sold or used as an independent product, it can be stored in a non-volatile computer-readable storage medium that is executable by a processor. Based on this understanding, the technical solution of the present application, or the part that contributes to the prior art or the part of the technical solution, can be embodied in the form of a software product, and the computer software product is stored in a storage medium, including a number of instructions to enable a computer device (which can be a personal computer, a server, or a network device, etc.) to perform all or part of the steps of the method described in each embodiment of the present application. The aforementioned storage medium includes: U disk, mobile hard disk, read-only memory (ROM), random access memory (RAM), disk or optical disk, etc., various media that can store program codes.

[0123] The above are only specific implementations of the present application, but the protection scope of the present application is not limited thereto. Any technician familiar with the technical field can easily think of changes or substitutions within the technical scope disclosed in the present application, which should be included in the protection scope of the present application. Therefore, the protection scope of the present application should be based on the protection scope of the claims.

Claims

1. A launch vehicle control method based on wind field characteristics, characterized in that: The method comprises: Obtaining a wind field model generated by fitting the actually measured wind field data; Determine the relative speed of the carrier rocket relative to the atmosphere in the aircraft coordinate system at the current moment according to the speed of the carrier rocket in the launch coordinate system at the current moment, the wind speed in the launch coordinate system at the current moment simulated in the wind field model, and a first conversion relationship, wherein the first conversion relationship is used to convert the launch coordinate system to the aircraft coordinate system; Calculating an influencing parameter through the relative speed, and calculating the aerodynamic force at the current moment according to the influencing parameter; Integrating the differential equations of the center of mass motion of the launch vehicle in combination with the aerodynamic force, the preset initial values ​​of the preset control parameters and the flight program angle expression of the launch vehicle that changes with time, calculating the preset state variables of the launch vehicle at the next moment, and thereby cyclically obtaining the preset state variables at each moment; Calculating the objective function of the launch vehicle based on the preset state variables at each moment; Based on the sequential quadratic programming algorithm, the value of the preset control parameter is changed through the objective function and the preset constraint condition, and the corresponding objective function under different preset control parameters is calculated to determine the optimized value of the preset control parameter.

2. The method according to claim 1, characterized in that The aerodynamic forces are calculated as follows: Calculate the influencing parameters and dynamic pressure affecting the aerodynamic coefficient by the speed of each coordinate axis of the aircraft coordinate system in the relative speed, wherein the influencing parameters include the angle of attack and the sideslip angle; The aerodynamic force is 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 by the influencing parameters.

3. The method according to claim 1, characterized in that The flight program angle expression is used to describe the program angle of the launch vehicle in each time period during the flight process, and the flight program angle expression includes: Among them, 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, and t1 ≤ t < t2 is the turning section of the launch vehicle. is the preset program angle, and 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, and t3 ≤ t < t4 is the constant 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.

4. The method according to claim 3, characterized in that The preset control parameters include the firing azimuth, the maximum 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 flight program angle slope, and the start time of the aiming section of the launch vehicle in the second stage flight phase, The time-varying angle of attack is described by the following formula: α(t)=-4×α M ×Z×(1-Z) Among them, α(t) is the angle of attack that changes with time, α M is the maximum absolute value of the angle of attack in the turning section, a is a preset constant value, t m It is the time when the angle of attack reaches the extreme value in the turning section.

5. The method according to claim 1, characterized in that The differential equations of the center of mass motion of the launch vehicle are described by the following formula: in, is the acceleration, is a second transformation relationship, wherein the second transformation relationship is used to transform the aircraft coordinate system into the launch coordinate system, is the apparent acceleration in the aircraft coordinate system, 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 time derivative of the position vector X, V is the velocity of the launch vehicle at each moment in the launch coordinate system, and g is the gravitational acceleration.

6. The method according to claim 1, characterized in that The objective function of the launch vehicle is calculated as follows: Calculating a target parameter value of the launch vehicle according to preset state variables at each moment, wherein the target parameter value includes at least one of the following: a payload under a given orbit entry condition, an elliptical orbit under a given payload, and a range; The objective function of the launch vehicle is calculated based on each target parameter value and its corresponding preset weight factor.

7. The method according to claim 1, characterized in that The optimized value of the preset control parameter is determined by: 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 length, updating the preset control parameter to obtain an updated value of the preset control parameter; Determine whether the preset convergence condition is met. If the preset convergence condition is not met, return to recalculate the preset control parameters of the launch vehicle until the preset convergence condition is met, and use the updated value of the preset control parameter corresponding to the convergence condition as the optimized value of the preset control parameter.

8. A launch vehicle control device based on wind field characteristics, characterized in that: The device comprises: An acquisition module is used to acquire a wind field model generated by fitting the actually measured wind field data; A determination module determines the relative speed of the carrier rocket relative to the atmosphere in the aircraft coordinate system at the current moment according to the speed of the carrier rocket in the launch coordinate system at the current moment, the wind speed in the launch coordinate system at the current moment simulated in the wind field model, and a first conversion relationship, wherein the first conversion relationship is used to convert the launch coordinate system to the aircraft coordinate system; A first calculation module calculates an influencing parameter through the relative speed, and calculates the aerodynamic force at the current moment according to the influencing parameter; The second calculation module integrates the differential equations of the center of mass 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 that changes with time, calculates the preset state variable of the launch vehicle at the next moment, and obtains the preset state variable at each moment in a cycle; A third calculation module calculates the objective function of the launch vehicle according to the preset state variables at each moment; The optimization module, based on the sequential quadratic programming algorithm, changes the value of the preset control parameter through the objective function and the preset constraint conditions, and calculates the corresponding objective function under different preset control parameters to determine the optimized value of the preset control parameter.

9. An electronic device, characterized in that: include: A processor, a memory and a bus, wherein the memory stores machine-readable instructions executable by the processor, and when the electronic device is running, the processor and the memory communicate through the bus, and the machine-readable instructions are executed by the processor to execute the steps of the launch vehicle control method based on wind field characteristics as described in any one of claims 1 to 7.

10. A computer-readable storage medium, characterized in that: The computer-readable storage medium stores a computer program, and when the computer program is executed by a processor, the steps of the launch vehicle control method based on wind field characteristics as described in any one of claims 1 to 7 are executed.

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