A launch vehicle load regulation and attitude control method and device based on radar wind measurement, medium and product

Abstract

The application discloses a launch vehicle load regulation and attitude control method and device based on radar wind measurement, a medium and a product, relates to the field of launch vehicle load reduction control, determines wind field information at a preset time based on one-dimensional wind speed measurement results of a laser radar, and obtains an attack angle at a current time, an engine swing angle at the current time and attitude information of a current position of the launch vehicle; according to the above data, the influence of the attack angle and the engine swing angle on the missile body bending moment of the launch vehicle is determined; according to the attitude information, a model predictive controller is solved to determine an optimized control quantity sequence; the model predictive controller comprises a system output equation and a cost function; according to the above data, an extended state observer is used to determine a disturbance compensation amount; according to the optimized control quantity sequence and the disturbance compensation amount, a final control quantity is determined to reduce the load of the launch vehicle. The application realizes accurate load reduction of the launch vehicle.

Description

TECHNICAL FIELD

[0001] The present application relates to the field of launch vehicle load control, and particularly relates to a launch vehicle load regulation and attitude control method, device, medium and product based on radar wind measurement. BACKGROUND

[0002] The main function of the launch vehicle control system is to ensure the stability and tracking of the attitude under various disturbances. When flying in the high dynamic pressure area, the aerodynamic load generated by the wind will affect the rocket. In order to maintain the stability and tracking of the attitude, the attitude control system must generate a control moment balanced with the aerodynamic load, which will make the rocket structure bear the bending moment generated by the interaction of the two, and thus the structural strength needs to be increased. With the increasing complexity of the rocket launch task, the requirements for structural efficiency and carrying capacity are continuously improved. In order to smoothly pass through the high dynamic pressure area under the limited structural strength index, active load alleviation control technology has been widely studied.

[0003] The idea of the existing active load alleviation control method is to alleviate the static bending moment load of the rocket structure. The common practice is to introduce feedback about the normal overload in the attitude control loop. This will bring the following problems: 1. Since the lateral acceleration information contains redundant information such as the normal component of the thrust force and the rotational inertia force in addition to the aerodynamic load, directly feeding back the acceleration will reduce the stability of the control system, and thus limit the load alleviation ratio; 2. For the control system, the load alleviation only concerns the characteristic quantity qα reflecting the load information, while the internal force bending moment actually affects the upper limit of the rocket structure bearing, and the bending moment upper limit borne by each section of the rocket body is inconsistent, so the structural load cannot be finely adjusted.

[0004] In summary, the existing active load alleviation control method cannot accurately alleviate the load. SUMMARY

[0005] The purpose of the present application is to provide a launch vehicle load regulation and attitude control method, device, medium and product based on radar wind measurement, to accurately alleviate the load of the launch vehicle.

[0006] To achieve the above purpose, the present application provides the following solutions:

[0007] In a first aspect, the present application provides a launch vehicle load regulation and attitude control method based on radar wind measurement, characterized in that it comprises:

[0008] obtaining a one-dimensional wind speed measurement result of a laser radar at a current time of a current position of a launch vehicle, an attack angle at the current time, an engine swing angle at the current time and attitude information at the current time; the one-dimensional wind speed measurement result is the projection of the wind speed measured by the laser radar in different measurement directions at a set position; the attitude information includes attitude angle and attitude angular velocity;

[0009] Determine wind field information of a preset time at a current position of the launch vehicle according to a one-dimensional wind speed measurement result of the laser radar at the current time; the wind field information includes a horizontal wind speed and a horizontal wind direction;

[0010] Determine an influence of the attack angle and the engine swing angle on a body bending moment of the launch vehicle according to the wind field information of the preset time, the attack angle at the current time and the engine swing angle at the current time; the body bending moment is determined according to a bending moment generated by aerodynamic force, a bending moment generated by control force and a bending moment generated by inertial force;

[0011] Solve a model predictive controller according to the attitude angle and the attitude angular velocity to determine an optimized control quantity sequence; the model predictive controller includes a system output equation and a cost function; the system output equation is a linear equation about state quantities, control quantities and disturbance quantities, and is determined according to the influence of the attack angle and the engine swing angle on the body bending moment of the launch vehicle; the state quantities are the attitude angle and the attitude angular velocity; the control quantity is the engine swing angle at the next time; the cost function is determined according to the system output equation;

[0012] Determine a disturbance compensation quantity by using an extended state observer according to the wind field information of the preset time, the attack angle at the current time, the engine swing angle at the current time, the attitude angle and the attitude angular velocity;

[0013] Determine a final control quantity according to the optimized control quantity sequence and the disturbance compensation quantity, and perform load reduction on the launch vehicle based on the final control quantity.

[0014] In a second aspect, a computer device is provided, which includes a memory, a processor, a computer program stored in the memory and executable on the processor, and the processor executes the computer program to implement the launch vehicle load regulation and attitude control method based on radar wind measurement according to any one of the above.

[0015] In a third aspect, a computer readable storage medium is provided, which stores a computer program, and the computer program is executed by a processor to implement the launch vehicle load regulation and attitude control method based on radar wind measurement according to any one of the above.

[0016] In a fourth aspect, a computer program product is provided, which includes a computer program, and the computer program is executed by a processor to implement the launch vehicle load regulation and attitude control method based on radar wind measurement according to any one of the above.

[0017] According to the specific embodiments provided in the present application, the following technical effects are disclosed:

[0018] The application provides a launch vehicle load regulation and attitude control method and device based on radar wind measurement, equipment, medium and product, which obtains wind field information at a preset moment of a current position of a launch vehicle, an attack angle at a current moment, an engine swing angle at the current moment and attitude information; the attitude information includes an attitude angle and an attitude angular velocity; influences of the attack angle and the engine swing angle on a missile body bending moment of the launch vehicle are determined according to the wind field information at the preset moment, the attack angle at the current moment and the engine swing angle at the current moment; a model predictive controller is solved according to the attitude angle and the attitude angular velocity, and an optimized control quantity sequence is determined; the model predictive controller includes a system output equation and a cost function; the system output equation is a linear equation about state quantities, control quantities and disturbance quantities, and is determined according to influences of the attack angle and the engine swing angle on the missile body bending moment of the launch vehicle; the cost function is determined according to the system output equation; a disturbance compensation quantity is determined by using an extended state observer according to the wind field information at the preset moment, the attack angle at the current moment, the engine swing angle at the current moment, the attitude angle and the attitude angular velocity; a final control quantity is determined according to the optimized control quantity sequence and the disturbance compensation quantity, and the launch vehicle is unloaded based on the final control quantity. The application realizes accurate unloading of the launch vehicle. BRIEF DESCRIPTION OF DRAWINGS

[0019] In order to more clearly illustrate the technical solutions in the embodiments of the present application or the prior art, the following will briefly introduce the drawings needed in the embodiments. Obviously, the drawings in the following description only constitute some embodiments of the present application, and for those skilled in the art, other drawings can also be obtained without creative labor.

[0020] Figure 1 A flowchart of a launch vehicle load regulation and attitude control method based on radar wind measurement provided by an embodiment of the present application;

[0021] Figure 2 A flowchart of the launch vehicle load regulation and attitude control method based on radar wind measurement of the present application in actual application;

[0022] Figure 3 A control block diagram of the launch vehicle load regulation and attitude control method based on radar wind measurement of the present application;

[0023] Figure 4 A schematic diagram of a rocket-borne radar;

[0024] Figure 5 A schematic diagram of a radar coordinate system, i.e. a measurement point position;

[0025] Figure 6 A schematic diagram of a rocket design load;

[0026] Figure 7 A curve diagram of unloading control effect;

[0027] Figure 8 for the attitude angle deviation curve due to load shedding;

[0028] Figure 9 for the model predictive load shedding control error and the dynamic pressure and angle of attack product considering disturbance compensation;

[0029] Figure 10 for the model predictive load shedding control error and qa;

[0030] Figure 11 for the structure schematic diagram of a computer device provided by an embodiment of the present application. DETAILED DESCRIPTION

[0031] The technical solutions in the embodiments of the present application will be described clearly and completely below with reference to the drawings in the embodiments of the present application. Obviously, the described embodiments are only part of the embodiments of the present application, rather than all the embodiments. Based on the embodiments in the present application, all other embodiments obtained by those skilled in the art without creative work fall within the scope of protection of the present application.

[0032] The above purposes, features and advantages of the present application can be more obvious and easy to understand. The present application will be described in further detail below with reference to the drawings and specific embodiments.

[0033] The present application models and forecasts future wind fields through laser radar measurement, and integrates future wind field information into model predictive control to optimize structural load and attitude tracking indicators. At the same time, an extended state observer is used to estimate and compensate disturbances, thereby improving overall control effect.

[0034] In an exemplary embodiment, as shown in Figures 1-3 a launch vehicle load regulation and attitude control method based on radar wind measurement is provided, including the following steps:

[0035] S1: obtaining one-dimensional wind speed measurement results of a laser radar at a current time at a current position of a launch vehicle, an angle of attack at a current time, an engine swing angle at a current time, and attitude information at a current time; the one-dimensional wind speed measurement results are projections of wind speeds measured by the laser radar at different measurement directions at a set position; the attitude information includes attitude angles and attitude angular velocities.

[0036] S2: determining wind field information at a preset time at a current position of the launch vehicle according to the one-dimensional wind speed measurement results of the laser radar at the current time; the wind field information includes horizontal wind speed and horizontal wind direction.

[0037] As an optional implementation, S2 specifically includes:

[0038] S21: Determine the projection of the wind speed at the preset time in the laser radar coordinate system according to the one-dimensional wind speed measurement result of the laser radar at the current time, combined with the position, speed and attitude information of the launch vehicle at the current time.

[0039] S22: According to the projection of the wind speed at the preset time in the laser radar coordinate system, combined with the coordinate conversion matrix from the geographic coordinate system to the radar coordinate system, determine the horizontal eastward wind speed and the horizontal northward wind speed in the geographic coordinate system at the preset time, and determine the horizontal wind direction to obtain the wind field information at the preset time.

[0040] In practical application, in view of the feature that the attitude of the launch vehicle changes constantly during flight, the application introduces a new group of radar measurement information, combines five groups of radar information, and deduces the inversion method of three-dimensional wind field considering the attitude of the rocket.

[0041] The steps of the wind field information estimation method based on the rocket-borne laser radar are as follows:

[0042] The wind speed can be represented by the movement speed of aerosol particles in the atmosphere. Based on the speed measurement principle of incoherent laser radar detection, the one-dimensional radial speed of aerosol particles in the atmosphere relative to the laser radar (i.e. the speed component of aerosol particles in the detection beam direction) can be obtained. By using a multi-beam measurement method, the one-dimensional radial speed in different directions can be measured, and the three-dimensional wind speed can be calculated from these one-dimensional radial speeds. The traditional wind speed measurement platform is usually fixed, while the rocket-borne laser radar is a high-speed variable attitude platform. In the wind speed calculation link, the motion characteristics of the rocket need to be considered, and the rocket-borne laser radar is shown in Figure 4 .

[0043] (1) Coordinate system definition and wind speed projection.

[0044] When the laser wind measurement radar works on the rocket, the radar system cannot remain horizontal like on the ground-based platform, and the influence of the attitude needs to be considered. It is assumed that the scanning angle of the radar at this time is still γ s , as shown in Figure 5 five directions to measure the radial wind speed, and the phase angle between adjacent directions is 90°. In order to describe the problem, first define the radar coordinate system S l (Ox l y l z l ) and its three-axis direction. The origin of the coordinate system is set at the center of the radar antenna, and the five measurement points 0, 1, 2, 3 and 4 are in the plane α; the Oz l axis passes through the origin and is perpendicular to the plane α, and the positive direction is consistent with the radar detection direction; define a plane β passing through the origin and parallel to the plane α, Ox lThe axis is in the plane β and parallel to the line connecting the measuring point 1 and the measuring point 3, and the positive direction is from the measuring point 3 to the measuring point 1; Oy l The axis is in the plane β and satisfies the right-hand rule.

[0045] The wind speed is often described in the local geographic coordinate system S d (Ox d y d z d ), so it is necessary to establish the transformation relationship between the two coordinate systems. A set of Euler angles can be used to describe the position relationship between S l and S d . Suppose that the geographic coordinate system S d is rotated around the z-axis, y-axis and x-axis in turn ψ l and γ l , and coincides with the radar coordinate system. Then the coordinate transformation matrix L d from the geographic coordinate system S l to the radar coordinate system S ld and the coordinate transformation matrix L l from the radar coordinate system S d to the geographic coordinate system S dl can be solved, as shown in formula (1) and formula (2):

[0046]

[0047] Where L x (γ l ) is the x-axis coordinate transformation matrix of the geographic coordinate system S d to the radar coordinate system S l ; L y (ψ l ) is the y-axis coordinate transformation matrix of the geographic coordinate system S d to the radar coordinate system S l ; is the z-axis coordinate transformation matrix of the geographic coordinate system S d to the radar coordinate system S l ; is the rotation angle of the geographic coordinate system S d rotating around the z-axis; ψ l is the rotation angle of the geographic coordinate system S d rotating around the y-axis; γ l is the rotation angle of the geographic coordinate system S d rotating around the x-axis; and T is the transpose.

[0048] The component column matrix of the wind speed vector in S d is:

[0049] (V wind )d = [u d v d w d ] T (3)

[0050] where u d , v d and w d are the projections of the wind velocity vector in the three axis directions of the geographical coordinate system S d .

[0051] Similarly, the components of the wind velocity vector in the radar coordinate system S l are:

[0052] (V wind ) l = [u l v l w l ] T (4)

[0053] where u l , v l and w l are the projections of the wind velocity vector in the three axis directions of the radar coordinate system S l .

[0054] The components of the wind velocity vector in the two coordinate systems satisfy the relationship:

[0055] (V wind ) d = L dl (V wind ) l (5)

[0056] (2) Three-dimensional wind field inversion algorithm considering attitude compensation.

[0057] During the flight of the rocket in the ascending phase, its attitude changes continuously, and the measurement results of the laser wind measurement radar directly installed on the rocket platform will be directly affected by the attitude changes. The present application considers this phenomenon and assumes that the wind field changes linearly, and derives the inversion method of three-dimensional wind field data.

[0058] When the wind field is uniformly distributed, the results at the four measurement directions (measurement points 1, 2, 3, and 4) are still the projections of the wind velocity vector in the laser emission direction, and at this time the measurement results are as follows: Figure 5

[0059]

[0060] where V1, V2, V3, and V4 represent the measurement results at the measurement points labeled 1, 2, 3, and 4, respectively. ​

[0061] The projection of the wind velocity vector in the radar coordinate system is solved:

[0062]

[0063] In practice, the wind field in the ascending stage of a launch vehicle is often described by the horizontal wind speed and the horizontal wind direction at a series of discrete height points. The wind speed and the wind direction vary linearly with height between adjacent discrete points, and the wind field is no longer uniform and stable.

[0064] The wind direction of high-altitude wind is relatively stable, and the horizontal wind direction also changes little with height. Assuming that the horizontal wind speed varies linearly with height and the horizontal wind direction remains unchanged, let Oz l The α plane intersects the measurement point 0, and the distance from the origin to the measurement point is L D , the altitude is h0, the horizontal wind speed is V H0 , and the horizontal wind direction is θ0. The horizontal wind speed and the horizontal wind direction satisfy the following rules:

[0065]

[0066] In the formula, h is the altitude of the measurement point, V H (h) and θ(h) are the horizontal wind speed and the horizontal wind direction at an altitude of h, respectively, and k w is the rate of change of the horizontal wind speed with height.

[0067] Take the measurement process of the measurement point 1 as an example. First, calculate the altitude of the measurement point 1. Let the position vector from the measurement point 0 to the measurement point 1 be r1, and according to the geometric relationship, we have:

[0068] (r1) l = [L D tanγ s 00] T (9)

[0069] In the formula, (r1) l is the position vector from the measurement point 0 to the measurement point 1 in the radar coordinate system.

[0070] According to formula (2), we have:

[0071]

[0072] In the formula, (r1) d is the position vector from the measurement point 0 to the measurement point 1 in the geographic coordinate system.

[0073] According to formula (10), the height difference between the measuring point 1 and the measuring point 0 can be solved. Similarly, the height differences between the measuring point 2, the measuring point 3 and the measuring point 4 and the measuring point 0 can be solved respectively.

[0074]

[0075] In the formula, Δh1 is the height difference between the measuring point 1 and the measuring point 0; Δh2 is the height difference between the measuring point 2 and the measuring point 0; Δh3 is the height difference between the measuring point 3 and the measuring point 0; and Δh4 is the height difference between the measuring point 4 and the measuring point 0.

[0076] The height difference of the measuring point 1 in formula (11) is substituted into formula (8) to obtain the wind field information at the measuring point 1.

[0077]

[0078] In the formula, V H1 is the horizontal wind field size at the measuring point 1; θ1 is the horizontal wind direction at the measuring point 1; (V 1w ) d is the projection of the wind speed at the measuring point 1 in the geographical coordinate system.

[0079] For the purpose of simplifying the expression, the elements in L ld are simply recorded in the form of a ij , as shown in formula (13).

[0080]

[0081] According to the coordinate transformation matrix L d from the geographical coordinate system S l to the radar coordinate system S ld , the theoretical measurement result expression of the radar at the four measuring points can be given:

[0082]

[0083] The wind speed vector at the measuring point 0 in the radar coordinate system S l is solved as:

[0084]

[0085] In the formula, V is the horizontal eastward component of the theoretically estimated wind speed in the radar coordinate system S l ; is the horizontal northward component of the theoretically estimated wind speed in the radar coordinate system S l ; is the vertical component of the theoretically estimated wind speed in the radar coordinate system S l . is the projection of the real wind speed vector at the preset moment in the laser radar coordinate system.

[0086] The component array of the real wind speed vector at point 0 in the radar coordinate system S l is:

[0087]

[0088] In the formula, (V0w)l is the projection of the real wind speed vector in the radar coordinate system S l ; (V0w)d is the projection of the real wind speed vector in the geographical coordinate system S d .

[0089] By comparing formula (15) and formula (16), it can be seen that in a linearly changing wind field, the previous solving method will bring a system error. Only when k w is equal to 0, that is, the wind field satisfies the assumption of uniform stability, can it be solved without error. Therefore, the observation result at a distance of L D along the positive direction of Oz l is introduced. Then the observation value V0 is:

[0090] V0 = V H0 (a 31 cosθ0 + a 32 sinθ0) (17)

[0091] Using the measurement results at the four original point positions, the equation can be constructed:

[0092]

[0093] In the formula:

[0094]

[0095] It can be proved that the trigonometric function coefficients in formula (17) and formula (19) are linearly independent, that is, the value of Pa 32 -Qa 31 is not equal to zero. The horizontal eastward wind speed u d and the horizontal northward wind speed v d at the preset moment can be solved:

[0096]

[0097] The horizontal wind speed and the horizontal wind direction at the preset moment can be further solved according to the eastward and northward wind speeds.

[0098] S2: determining the influence of the angle of attack and the engine swing angle on the rocket body bending moment of the launch vehicle according to the preset time wind field information, the current time angle of attack and the current time engine swing angle; the rocket body bending moment is determined according to the bending moment generated by the aerodynamic force, the bending moment generated by the control force and the bending moment generated by the inertial force.

[0099] In practical applications, a rocket design load diagram is as shown in Figure 6 When the bending moment at a certain position of the rocket body is considered as the output of the control system, it can be regarded as a function of the angle of attack and the engine swing angle, and the influences of the two on the rocket body bending moment are considered respectively. When calculating the bending moment, the influences of the aerodynamic force, the control force and the inertial force can be considered respectively.

[0100] As an optional implementation, the rocket body bending moment is determined according to the bending moment generated by the aerodynamic force, the bending moment generated by the control force and the bending moment generated by the inertial force, and specifically includes:

[0101] (1) Bending moment load generated by aerodynamic force.

[0102] The case of a rocket with a total length of L in the nominal flight is considered. According to the overall design and aerodynamic calculation, the aerodynamic lift distribution function of the rocket at this time is defined as:

[0103] F ayb (α, x, Ma), 0≤x≤L (21)

[0104] In the formula, F ayb represents the aerodynamic force function of the system, α is the current time angle of attack, x represents the integral variable, and Ma represents the Mach number.

[0105] The bending moment (the bending moment generated by the aerodynamic force) Ma(xk) generated at the position x k in the geometric coordinate system is:

[0106]

[0107] (2) Bending moment load generated by control force.

[0108] The bending moment (the bending moment generated by the control force) Mc(xk) generated by the control force at the position x k is:

[0109]

[0110] Wherein, x c is the position of the control force; F cyb is the component of the thrust in the y-axis of the system; and L is the length of the launch vehicle.

[0111] (3) Bending moment load generated by inertial force.

[0112] Since the rocket is not in force balance during flight, there are acceleration and angular acceleration. The inertial force introduced to ensure balance will also produce bending moment. The acceleration ay(xk) at position x k

[0113]

[0114] wherein a yb is the translational acceleration of the center of mass, is the rotational acceleration around the center of mass, x g represents the position of the center of mass. The bending moment M i (x k ) generated by the inertial force is:

[0115]

[0116] The sum of the bending moment generated by the aerodynamic force, the bending moment generated by the control force, and the bending moment generated by the inertial force is calculated to obtain the rocket body bending moment.

[0117] In practical applications, according to formula (22), formula (23), and formula (24), the total bending moment (rocket body bending moment) M(x k ) expression is as follows:

[0118] M(x k ) = M a (x k ) + M c (x k ) + M i (x k )(26)

[0119] During the flight of the rocket, the aerodynamic force and the control force are factors affecting the bending moment, and the small perturbation equation of the bending moment with respect to the angle of attack and the engine swing angle can be written as:

[0120]

[0121] wherein M′ α (x k ) is the influence of the angle of attack on the rocket body bending moment of the carrier rocket; is the influence of the engine swing angle on the rocket body bending moment of the carrier rocket; Δα is the angle of attack change; α w is the angle of attack at a preset time; is the engine swing angle change.

[0122]

[0123] is the nominal normal aerodynamic force coefficient; q is the dynamic pressure; S​m S is the reference area of the missile body; F ay F is the normal total aerodynamic force of the launch vehicle; m'(x) is the mass distribution function of the missile body; m is the total mass of the missile body; M azb M is the pitch channel aerodynamic moment; J z J is the z-axis rotational inertia; K is the normal component coefficient of the swing jet; F cyb F is the component of the thrust in the y-axis of the system; M czb M is the pitch channel control moment. The system is a right-handed rectangular coordinate system; the origin of the right-handed rectangular coordinate system is the center of mass O of the missile body, the Ox axis is parallel to the missile body axis and points to the head, the Oy axis is in the longitudinal symmetry plane of the missile body, the Oy axis is perpendicular to the Ox axis and points upward, and the Oz axis is perpendicular to the Oxy plane and points to the right.

[0124] As an optional implementation, S2 specifically comprises:

[0125] S21: determining the attack angle at the preset moment and the engine swing angle at the preset moment according to the wind field information at the preset moment.

[0126] S22: determining the attack angle change according to the attack angle at the preset moment and the attack angle at the current moment.

[0127] S23: determining the engine swing angle change according to the engine swing angle at the preset moment and the engine swing angle at the current moment.

[0128] S24: determining the influence of the attack angle on the bending moment of the missile body of the launch vehicle and the influence of the engine swing angle on the bending moment of the missile body of the launch vehicle.

[0129] S25: determining the influence of the attack angle and the engine swing angle on the bending moment of the missile body of the launch vehicle according to the attack angle at the preset moment, the attack angle change, the engine swing angle change, the influence of the attack angle on the bending moment of the missile body of the launch vehicle, and the influence of the engine swing angle on the bending moment of the missile body of the launch vehicle.

[0130] S3: solving the model predictive controller according to the attitude angle and the attitude angular velocity to determine the optimized control quantity sequence; the model predictive controller comprises a system output equation and a cost function; the system output equation is a linear equation about state quantities, control quantities and disturbance quantities, which is determined according to the influence of the attack angle and the engine swing angle on the bending moment of the missile body of the launch vehicle; the state quantities are the attitude angle and the attitude angular velocity; the control quantity is the engine swing angle at the next moment; and the cost function is determined according to the system output equation.

[0131] In practical applications, model predictive control is first applied in the field of industrial control. Thanks to its ability to solve multi-constrained optimization problems, model predictive control has shown broad prospects for development in many fields, including spacecraft control tasks. The essence of rolling model predictive control is to solve the open-loop optimization problem of the control sequence, and then apply the first element of the optimization sequence result to the controlled system to complete the closed loop. Model predictive control needs to establish a prediction model first, and then design a cost function to convert the control task into solving an optimization problem.

[0132] (1) The establishment of the prediction model (system output equation).

[0133] The small perturbation state equation of the launch vehicle is as follows:

[0134]

[0135] In the formula:

[0136]

[0137] Where, A c is the state matrix; B dc is the disturbance matrix; B uc is the control matrix; b1 is the influence coefficient of angular velocity on angular acceleration; b2 is the influence coefficient of attack angle on angular acceleration; b3 is the influence coefficient of engine swing spray angle on angular acceleration; c1 is the influence coefficient of attack angle on trajectory inclination angle change rate; c1' is the influence coefficient of attack angle on trajectory inclination angle change rate caused by wind; c2 is the influence coefficient of trajectory inclination angle on trajectory inclination angle change rate; c3 is the influence coefficient of engine swing spray angle on trajectory inclination angle change rate.

[0138] Discretization of the above state equation (30) can obtain the discrete system:

[0139] x(k+1) = Ax(k) + B d d(k) + B u u(k) (31)

[0140] In the formula:

[0141]

[0142] Based on the above discrete system, the predicted values of the state and output of the system at the next N time points can be given

[0143]

[0144] In the formula, U k is the control sequence; D kΦ0is the state transition matrix from time k to time k+N-1, Φ1is the state transition matrix from time k+1 to time k+N, Γ u0 Φ0is the state transition matrix from time k to time k+N-1, Φ1is the state transition matrix from time k+1 to time k+N, Γ u1 Φ0is the state transition matrix from time k to time k+N-1, Φ1is the state transition matrix from time k+1 to time k+N, Γ d0 Φ0is the state transition matrix from time k to time k+N-1, Φ1is the state transition matrix from time k+1 to time k+N, Γ d1 Φ0is the state transition matrix from time k to time k+N-1, Φ1is the state transition matrix from time k+1 to time k+N, Γ

[0145]

[0146] The output equation of the system is also considered. In general, the load reduction task is to reduce the total attack angle of the rocket during flight, that is, to reduce the load on the rocket body. More essentially, if the output of the system can directly reflect the bending moment at a certain place of the rocket body, better load reduction effect can be achieved. With the nominal trajectory, the attack angle deviation of the rocket and the bending moment deviation at a certain place of the rocket body can also be expressed as a linear combination of state quantities, disturbance quantities and control quantities. In order to unify the conclusions, the system output equation is:

[0147] y(k)=Cx(k)+D u u(k)+D d d(k)(37)

[0148] In the formula: y(k) is the influence of the attack angle and the engine swing angle on the bending moment of the rocket body of the launch vehicle; x(k) is the state quantity; u(k) is the control quantity; d(k) is the disturbance quantity. C is the state output matrix; D u is the control output matrix; D d is the disturbance output matrix.

[0149]

[0150] (2) Cost function design and optimization problem solving.

[0151] In the load reduction control task, the model prediction needs to consider the attitude tracking accuracy of the rocket, the change of the control quantity and the system output. Therefore, the cost function can be designed as:

[0152] J=J1+J2+J3

[0153]

[0154] Wherein, J1is the attitude tracking accuracy index; J2is the control quantity index; J3is the output index; x(k+N) is the state at time k+N; x r(k+N) is the expected state at time k+N; S is the state performance index matrix; u(k+i) is the control quantity at time k+i; u r (k+i) is the expected control quantity at time k+i; R i is the control performance index matrix; y(k+i) is the output at time k+i; y r (k+i) is the expected output at time k+i; Q i is the output performance indicator matrix.

[0155] Expand formula (39):

[0156] J1=J 11 +J 12 +J 13

[0157]

[0158] Among them, the part that can be used for optimization is only J 11 With J 12 Similarly, for J2 and J3:

[0159]

[0160] Further simplifying, the model predictive control problem is transformed into an optimal control problem. First consider J 11 ,J 12 ,J 13 sum:

[0161]

[0162] Where: X k+1 is the state prediction sequence from time k+1 to time k+N; is the transpose of the state prediction sequence from time k+1 to time k+N; Ω, Ψ, Θ1, Θ2, Θ3 are intermediate variables; (T in the upper right corner indicates transposition); D k is the disturbance sequence from time k to time k+N-1; (The T in the upper right corner indicates transposition).

[0163]

[0164] Among them, Q N-1 Output performance indicator matrix for the k+N-1th moment; R N-1 is the control performance indicator matrix at the k+N-1th moment.

[0165] J sum1 Some constant terms that do not have an optimized design space are discarded in the expression of . The constant term, which cannot be optimized, is also included in the middle term, and is kept for the sake of the neatness of the expression.

[0166] Substitute equation (33) into equation (42), we have:

[0167]

[0168] Where: H, F, E1 are intermediate variables.

[0169]

[0170] Similarly, consider the part of J2 that can be optimized J sum2 :

[0171]

[0172] Where: Γ u ′1, Ψ′, U rk , Ω1′, Y rk , Ω2′, E2 are intermediate variables.

[0173] Γ u ′1= [A N-1 B u A N-2 B u …B u ](47)

[0174]

[0175] According to the component index expressions, i.e. equation (42) and equation (46), the total index is:

[0176]

[0177] E = E1 + E2 (55)

[0178] Solve the partial derivative expression of J with respect to U k , and set the partial derivative to be a zero vector, then the optimized control quantity sequence at this time can be solved:

[0179]

[0180] Select the first element u * (k) in the sequence as the expected input value of the current system. At the next sampling time, repeat the above process.

[0181] S4: According to the preset time wind field information, the current time attack angle, the current time engine swing angle, the attitude angle and the attitude angle velocity, a disturbance compensation quantity is determined by using an extended state observer.

[0182] As an optional implementation, S4 specifically comprises:

[0183] According to the preset time wind field information, a preset time attack angle and a preset time engine swing angle are determined.

[0184] According to the preset time attack angle and the current time attack angle, an attack angle change amount is determined.

[0185] According to the preset time engine swing angle and the current time engine swing angle, an engine swing angle change amount is determined.

[0186] In actual application, in actual flight, there is an error between the real coefficient and the nominal coefficient, and in serious cases, the model mismatching occurs. In addition, there may be interference outside the model. These will have an adverse effect on the load reduction control. A suitable extended state observer can be designed to compensate for the model error and external interference.

[0187]

[0188] In the formula, is the attitude angle; is the attitude angular velocity; is the attitude angular acceleration; is the dynamic small disturbance coefficient, and Δα is the attack angle change amount; α w is the preset time attack angle; is the engine swing angle change amount; M d is the model parameter uncertainty, external disturbance and measurement uncertainty, and the state quantity and the control quantity are:

[0189]

[0190] Wherein, x1 is the attitude angle; x2 is the attitude angular velocity; x3 is the total disturbance; u is the engine swing angle change amount.

[0191] According to the preset time attack angle, the attack angle change amount, the engine swing angle change amount, the attitude angle and the attitude angular velocity, the state function of the launch vehicle is determined, that is, the state function of the system can be represented as:

[0192]

[0193] Wherein, is the first derivative of the attitude angle; is the first derivative of the attitude angular velocity; is the first derivative of the attitude angular acceleration; f is the derivative of the total disturbance.

[0194] According to the state function, an expansion state observer of the carrier rocket is determined, that is, an expansion state observer designed according to the state function is:

[0195]

[0196] Wherein, e1 is the pitch angle estimation error; z1 is the pitch angle estimation value; is the differential of the pitch angle estimation value; z2 is the pitch angle rate estimation value; is the differential of the pitch angle rate; z3 is the estimation of the disturbance; β1 is the first observer gain; β2 is the second observer gain; and β3 is the third observer gain.

[0197] The expansion state observer is solved to determine the disturbance compensation quantity.

[0198] By reasonably selecting the coefficients β1, β2 and β3, the following can be achieved:

[0199]

[0200] The disturbance estimation rudder deflection is:

[0201]

[0202] S5: According to the optimized control quantity sequence and the disturbance compensation quantity, a final control quantity is determined, and the carrier rocket is unloaded based on the final control quantity. In actual application, the final control quantity is the sum of the optimized control quantity sequence and the disturbance compensation quantity.

[0203] The application provides a carrier rocket load regulation and attitude control method based on radar wind measurement. The future wind field is measured and predicted by a laser radar, the structural load is optimized by using model predictive control, and the disturbance is estimated by using an expansion state observer, so that the control effect is improved. For an ideal unloading controller, on the one hand, it needs to have good unloading effect, and on the other hand, it needs to have certain anti-interference ability.

[0204] The application takes a digital simulation case as a method demonstration. The simulation working condition selects an atmospheric layer rising segment flight task of a certain type of carrier rocket. By using the carrier rocket load regulation and attitude control method based on radar wind measurement provided by the application, the unloading control effect is as shown in Figure 7 , the attitude angle deviation caused by unloading is as shown in Figure 8 , and the anti-interference ability test simulation result is as shown in Figure 9 and Figure 10 .

[0205] In an exemplary embodiment, a computer device can be provided, which can be a server or a terminal, and the internal structure diagram thereof can be as shown in Figure 11As shown in the figure. The computer device includes a processor, a memory, an input / output interface (Input / Output, referred to as I / O) and a communication interface. Among them, the processor, the memory and the input / output interface are connected through the system bus, and the communication interface is connected to the system bus through the input / output interface. Among them, the processor of the computer device is used to provide computing and control capability. The memory of the computer device includes a non-volatile storage medium and an internal memory. The non-volatile storage medium stores an operating system, a computer program and a database. The internal memory provides an environment for the operation of the operating system and the computer program in the non-volatile storage medium. The database of the computer device is used to store the attitude information, the attack angle and the engine swing angle data of the launch vehicle. The input / output interface of the computer device is used to exchange information between the processor and the external device. The communication interface of the computer device is used to communicate with the terminal outside through the network connection. The computer program is executed by the processor to realize a launch vehicle load regulation and attitude control method based on radar wind measurement.

[0206] Those skilled in the art can understand that, Figure 11 The structure shown in the figure is only a block diagram of part of the structure related to the scheme of the present application, and does not constitute a limitation on the computer device to which the scheme of the present application is applied. The specific computer device can include more or fewer components than those shown in the figure, or combine certain components, or have a different component arrangement. In one exemplary embodiment, a computer device is provided, including a memory and a processor, the memory storing a computer program, and the processor executing the computer program to realize the above-mentioned launch vehicle load regulation and attitude control method based on radar wind measurement.

[0207] In one exemplary embodiment, a computer readable storage medium is provided, storing a computer program, which is executed by a processor to realize the above-mentioned launch vehicle load regulation and attitude control method based on radar wind measurement.

[0208] In one exemplary embodiment, a computer program product is provided, including a computer program, which is executed by a processor to realize the above-mentioned launch vehicle load regulation and attitude control method based on radar wind measurement.

[0209] It should be noted that the user information (including but not limited to user equipment information, user personal information, etc.) and data (including but not limited to data for analysis, stored data, displayed data, etc.) involved in the present application are all information and data authorized by the user or authorized by all parties, and the collection, use and processing of related data need to comply with relevant regulations.

[0210] Those skilled in the art can understand that all or part of the processes in the above-mentioned embodiment methods can be completed by instructing the relevant hardware through a computer program. The computer program can be stored in a non-volatile computer readable storage medium, and when the computer program is executed, the processes of the above-mentioned embodiments of the methods can be included. Any reference to memory, database or other medium used in the embodiments provided in the present application can include at least one of non-volatile and volatile memory. Non-volatile memory can include read-only memory (ROM), magnetic tape, floppy disk, flash memory, optical storage, high-density embedded non-volatile memory, resistive memory (ReRAM), magnetoresistive random access memory (MRAM), ferroelectric memory (FRAM), phase change memory (PCM), graphene memory, etc. Volatile memory can include random access memory (RAM) or external cache memory, etc. As an illustration but not limitation, RAM can be in various forms, such as static random access memory (SRAM) or dynamic random access memory (DRAM), etc.

[0211] The database involved in the embodiments provided in the present application can include at least one of a relational database and a non-relational database. The non-relational database can include a distributed database based on a blockchain, etc., without being limited thereto. The processor involved in the embodiments provided in the present application can be a general-purpose processor, a central processing unit, a graphics processing unit, a digital signal processor, a programmable logic device, a data processing logic device based on quantum computing, etc., without being limited thereto.

[0212] The technical features of the above embodiments can be combined arbitrarily. In order to make the description concise, all possible combinations of the technical features in the above embodiments are not described, however, as long as the combinations of the technical features do not exist contradictory, they should be considered as the scope of the present application.

[0213] The principles and implementation modes of the present application are described by applying specific examples in the present application. The above-mentioned embodiments are only used to help understand the method and its core idea of the present application; at the same time, for those skilled in the art, according to the idea of the present application, the specific implementation mode and application range will be changed. In conclusion, the content of the present application should not be understood as a limitation.

Claims

1. A method for load regulation and attitude control of a launch vehicle based on radar wind measurement, characterized in that: include: Obtain the one-dimensional wind speed measurement result, the current angle of attack, the current engine swing angle and the current attitude information of the current position of the launch vehicle by the lidar at the current moment; The one-dimensional wind speed measurement result is the projection of the wind speed measured by the laser radar in different measurement directions at the set position; the attitude information includes the attitude angle and the attitude angular velocity; Determine the wind field information at the current position of the launch vehicle at a preset time based on the one-dimensional wind speed measurement result of the laser radar at the current time; the wind field information includes horizontal wind speed and horizontal wind direction; Determining, based on the wind field information at the preset time, the angle of attack at the current time, and the engine swing angle at the current time, the influence of the angle of attack and the engine swing angle on the body bending moment of the carrier rocket; The arrow body bending moment is determined based on the bending moment generated by aerodynamic force, the bending moment generated by control force and the bending moment generated by inertial force; Solving a model predictive controller based on the attitude angle and the attitude angular velocity to determine an optimized control variable sequence; the model predictive controller includes a system output equation and a cost function; the system output equation is a linear equation involving state variables, control variables, and interference variables determined based on the influence of the angle of attack and the engine swing angle on the bending moment of the launch vehicle; the state variables are the attitude angle and the attitude angular velocity; The control variable is the engine swing angle at the next moment; The cost function is determined based on the system output equation; Determining an interference compensation amount using an extended state observer based on the wind field information at the preset time, the angle of attack at the current time, the engine swing angle at the current time, the attitude angle, and the attitude angular velocity; A final control quantity is determined according to the optimized control quantity sequence and the interference compensation quantity, and the carrier rocket is deloaded based on the final control quantity.

2. The method for load regulation and attitude control of a launch vehicle based on radar wind measurement according to claim 1, characterized in that: Based on the one-dimensional wind speed measurement results of the lidar at the current moment, the wind field information at the preset moment of the current position of the carrier rocket is determined, specifically including: Based on the one-dimensional wind speed measurement results of the lidar at the current moment, combined with the position, speed and attitude information of the carrier rocket at the current moment, the projection of the wind speed at the preset moment in the lidar coordinate system is determined; According to the projection of the wind speed at the preset time in the lidar coordinate system, combined with the coordinate conversion matrix from the geographic coordinate system to the radar coordinate system, the horizontal eastward wind speed and the horizontal northward wind speed in the geographic coordinate system at the preset time are determined, and the horizontal wind direction is determined to obtain the wind field information at the preset time.

3. The method for load regulation and attitude control of a launch vehicle based on radar wind measurement according to claim 1, characterized in that: Determining, based on the wind field information at the preset time, the angle of attack at the current time, and the engine swing angle at the current time, an influence of the angle of attack and the engine swing angle on the body bending moment of the carrier rocket specifically includes: Determining the angle of attack and the engine swing angle at the preset time according to the wind field information at the preset time; Determining a change in the angle of attack according to the angle of attack at the preset time and the angle of attack at the current time; determining an engine swing angle change according to the engine swing angle at the preset time and the engine swing angle at the current time; Determining the influence of the angle of attack on the bending moment of the rocket body of the launch vehicle and the influence of the engine swing angle on the bending moment of the rocket body of the launch vehicle; According to the angle of attack at the preset moment, the change in the angle of attack, the change in the engine swing angle, the influence of the angle of attack on the bending moment of the rocket body of the carrier rocket, and the influence of the engine swing angle on the bending moment of the rocket body of the carrier rocket, the influence of the angle of attack and the engine swing angle on the bending moment of the rocket body of the carrier rocket is determined.

4. The method for load regulation and attitude control of a launch vehicle based on radar wind measurement according to claim 3, characterized in that: Determining the influence of the angle of attack on the body bending moment of the launch vehicle and the influence of the engine swing angle on the body bending moment of the launch vehicle specifically includes: Using the formula Determining the influence of the angle of attack on the bending moment of the rocket body of the launch vehicle and the influence of the engine swing angle on the bending moment of the rocket body of the launch vehicle; Among them, M′ α (x k ) is the effect of the angle of attack on the bending moment of the launch vehicle; The effect of the engine swing angle on the rocket body bending moment of the carrier rocket; is the nominal normal aerodynamic coefficient; q is the dynamic pressure; S m is the reference area of ​​the arrow body; F ay is the total aerodynamic force in the normal direction of the launch vehicle; x k The position for calculating bending moment; F ayb is the aerodynamic function of the system; x is the integral variable; m′(x) is the mass distribution function of the rocket body; m is the total mass of the rocket body; M azb is the aerodynamic moment of the pitch channel; J z is the moment of inertia along the z axis; x g is the center of mass position; K is the normal component coefficient of the pendulum jet; x c is the position where the control force acts; F cyb is the component of thrust on the y-axis of the system; M czb is the pitch channel control torque; L is the length of the launch vehicle; this system is a right-handed rectangular coordinate system; the origin of the right-handed rectangular coordinate system is the center of mass O of the rocket body, the Ox-axis is parallel to the axis of the rocket body and points to the head, the Oy-axis is in the longitudinal symmetry plane of the rocket body, the Oy-axis is perpendicular to the Ox-axis and points upward, and the Oz-axis is perpendicular to the Oxy plane and points to the right.

5. The method for load regulation and attitude control of a launch vehicle based on radar wind measurement according to claim 1, characterized in that: The bending moment of the rocket body is determined based on the bending moment generated by aerodynamic force, control force and inertia force, specifically including: Using the formula Determine the bending moment caused by aerodynamic forces; where M a (x k ) is the bending moment caused by aerodynamic force; x k The position for calculating bending moment; F ayb is the aerodynamic function of the system, α is the angle of attack at the current moment, x represents the integral variable, and Ma represents the Mach number; Using the formula Determine the bending moment generated by the control force; where M c (x k ) is the bending moment generated by the control force; x c is the position where the control force acts; F cyb is the component of thrust on the y-axis of this system; L is the length of the launch vehicle; Using the formula Determine the bending moment caused by the inertial force; where M i (x k ) is the bending moment generated by the inertial force; m′(x) is the mass distribution function of the rocket body; a yb (x) is the translational acceleration of the center of mass; The sum of the bending moment generated by the aerodynamic force, the bending moment generated by the control force, and the bending moment generated by the inertial force is calculated to obtain the arrow body bending moment.

6. The method for carrier rocket payload regulation and attitude control based on radar wind measurement according to claim 1, characterized in that: The system output equation is: y(k)=Cx(k)+D u u(k)+D d d(k); Wherein, y(k) is the influence of the angle of attack and the engine swing angle on the bending moment of the rocket body; x(k) is the state quantity; u(k) is the control quantity; d(k) is the interference quantity; C, D u 、D d is the intermediate variable; M′ α (x k ) is the effect of the angle of attack on the bending moment of the launch vehicle; is the influence of the engine swing angle on the bending moment of the rocket body of the carrier rocket.

7. The method for load regulation and attitude control of a launch vehicle based on radar wind measurement according to claim 1, characterized in that: Determining the interference compensation amount using an extended state observer based on the wind field information at the preset time, the angle of attack at the current time, the engine swing angle at the current time, the attitude angle, and the attitude angular velocity specifically includes: Determining the angle of attack and the engine swing angle at the preset time according to the wind field information at the preset time; Determining a change in the angle of attack according to the angle of attack at the preset time and the angle of attack at the current time; determining an engine swing angle change according to the engine swing angle at the preset time and the engine swing angle at the current time; The state function of the carrier rocket is determined according to the angle of attack at the preset time, the change in the angle of attack, the change in the engine swing angle, the attitude angle, and the attitude angular velocity; the state function is: in, is the first-order derivative of the attitude angle; x2 is the attitude angular velocity; is the first-order derivative of attitude angular velocity; x3 is attitude angular acceleration; is the first derivative of attitude angular acceleration; is the dynamic small perturbation coefficient; is the attitude angle; Δα is the change in angle of attack; α w is the angle of attack at the preset moment; u is the change in engine swing angle; f is the derivative of the total disturbance; An extended state observer of the launch vehicle is determined according to the state function; the extended state observer of the launch vehicle is: Among them, e1 is the pitch angle estimation error; z1 is the pitch angle estimation value; is the differential of the estimated pitch angle; z2 is the estimated pitch angle rate; is the differential of the pitch angular rate; z3 is the estimate of the disturbance; β1 is the first observer gain; β2 is the second observer gain; β3 is the third observer gain; The extended state observer is solved to determine the disturbance compensation amount.

8. A computer device comprising: A memory and a processor, wherein the computer program is stored in the memory and can be run on the processor, and the processor executes the computer program to implement the carrier rocket payload regulation and attitude control method based on radar wind measurement according to any one of claims 1 to 7.

9. A computer-readable storage medium having a computer program stored thereon, characterized in that: When the computer program is executed by a processor, the method for load regulation and attitude control of a carrier rocket based on radar wind measurement according to any one of claims 1 to 7 is implemented.

10. A computer program product comprising a computer program, characterized in that When the computer program is executed by a processor, the method for load regulation and attitude control of a carrier rocket based on radar wind measurement according to any one of claims 1 to 7 is implemented.

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