A method, system and device for controlling the ascent phase of a launch vehicle

Abstract

The present invention discloses a control method, system and equipment for the ascent stage of a carrier rocket, and relates to the field of rocket flight control. A rigid body dynamics model of the ascent stage of a carrier rocket and an elastic vibration model of the ascent stage of a carrier rocket are constructed; a dynamics model of the ascent stage of a carrier rocket is formed; the model is rewritten into a velocity coordinate system and linearized to obtain a small perturbation linearization model; an extended state observer of a rocket system is obtained according to the small perturbation linearization model; the stability of the extended state observer is derived to obtain a first ESO open-loop transfer function and a second ESO open-loop transfer function; the first preset parameter is adjusted to obtain a first control result of the ascent stage of a carrier rocket; the second preset parameter is adjusted to obtain a second control result of the ascent stage of a carrier rocket; the first control result of the ascent stage of a carrier rocket and the second control result of the ascent stage of a carrier rocket represent the control result of the ascent stage of a carrier rocket. The present invention improves the accuracy of attitude stability tracking and load reduction control during the flight of the ascent stage of a carrier rocket.

Description

Technical Field

[0001] The present invention relates to the field of rocket flight control technology, and in particular to a control method, system and equipment for the ascent phase of a carrier rocket. Background Art

[0002] During the ascent phase of the launch vehicle's flight, the control target is not unique, but interference still exists. The existence of this part of interference will also affect the load reduction control task, and it is necessary to study the interference suppression method under multiple control targets.

[0003] In the traditional launch vehicle controller design with PID control as the core, the elimination of steady-state errors and the suppression of interference are achieved through the integral feedback of attitude angle deviation. However, during the load reduction control process, the attitude angle needs to be changed to adapt to the changing wind attack angle, which will cause attitude angle deviation. Therefore, during the load reduction process, appropriate attitude angle deviation is allowed. However, the integral link cannot distinguish and will offset this part of the attitude angle error as interference, thereby reducing the load reduction effect. Summary of the invention

[0004] The purpose of the embodiments of the present invention is to provide a method, system and device for controlling the ascent phase of a launch vehicle, which improves the accuracy of attitude stability tracking and load reduction control during the flight of the ascent phase of the launch vehicle.

[0005] To achieve the above purpose, the embodiment of the present invention provides the following solution:

[0006] A method for controlling an ascent phase of a launch vehicle, comprising:

[0007] Acquire rigid body dynamics data and elastic vibration data of the ascent stage of the launch vehicle; construct a rigid body dynamics model of the ascent stage of the launch vehicle according to the rigid body dynamics data of the ascent stage; and construct an elastic vibration model of the ascent stage of the launch vehicle according to the elastic vibration data of the ascent stage;

[0008] The rigid body dynamics model of the launch vehicle's ascent stage and the elastic vibration model of the launch vehicle's ascent stage constitute a dynamics model of the launch vehicle's ascent stage; the dynamics model of the launch vehicle's ascent stage is rewritten into a velocity coordinate system and linearized to obtain a small disturbance linearization model;

[0009] An extended state observer of the rocket system is obtained according to the small disturbance linearization model; the working modes of the extended state observer include: an attitude tracking control mode and / or a load reduction control mode; stability derivation is performed on the extended state observer in the attitude tracking control mode to obtain a first ESO open-loop transfer function; stability derivation is performed on the extended state observer in the load reduction control mode to obtain a second ESO open-loop transfer function;

[0010] According to the first ESO open-loop transfer function and the feedback control law, the extended state observer in the attitude tracking control mode is adjusted by the first preset parameters to obtain the first control result of the launch vehicle ascent phase; the first preset parameters include: controller frequency and controller damping ratio;

[0011] According to the second ESO open-loop transfer function and the feedback control law, the second preset parameter adjustment is performed on the extended state observer under the load reduction control mode to obtain the second control result of the ascent phase of the launch vehicle; the second preset parameters include: controller frequency, controller damping ratio and overload feedback coefficient; the first control result of the ascent phase of the launch vehicle and the second control result of the ascent phase of the launch vehicle are used to characterize the control result of the ascent phase of the launch vehicle.

[0012] Optionally, constructing a rigid body dynamics model of the ascent stage of a launch vehicle according to the rigid body dynamics data of the ascent stage specifically includes:

[0013] ;

[0014] Where: is the launch system position vector; is the launch system velocity vector; For quality; is the angular velocity vector, are the three-axis angular velocities respectively; is the three-axis attitude angle; is the thrust vector, is the thrust torque vector; is the aerodynamic force vector, is the aerodynamic moment vector; is the gravitational acceleration vector; is the Coriolis acceleration vector; is the implicate acceleration vector; is the function of the consumption in seconds over time; is the inertia matrix;

[0015] The elastic vibration model of the ascent stage of the launch vehicle is constructed according to the elastic vibration data of the ascent stage, specifically comprising: ;

[0016] Where: For the pitch channel Generalized coordinates of elastic vibration of order modes; They are pitch channel Natural frequencies and damping coefficients of order modes; They represent the pitch channel engine swing angle and angular acceleration respectively; and They are the pitch channel engine swing angle and angular acceleration The influence coefficient of the order vibration mode; the rigid body dynamics model of the launch vehicle ascent stage and the elastic vibration model of the launch vehicle ascent stage constitute the dynamics model of the launch vehicle ascent stage, which specifically includes: the elastic vibration is coupled with the dynamics model of the launch vehicle ascent stage by affecting the attitude angle and angular velocity sensed by the inertial group.

[0017] Optionally, the dynamic model of the launch vehicle's ascent phase is rewritten into a velocity coordinate system and linearized to obtain a small-disturbance linearized model, specifically including: rewriting the dynamic model of the launch vehicle's ascent phase into a velocity coordinate system, performing control algorithm design and closed-loop system stability analysis and linearizing to obtain the small-disturbance linearized model.

[0018] Optionally, obtaining an extended state observer of a rocket system according to the small disturbance linearization model specifically includes:

[0019] ;

[0020] Where: is the small disturbance coefficient of the control quantity under the nominal state; is the observer parameter, by setting the observer parameter reasonably; , ; The attitude tracking control mode of the extended state observer specifically includes: adopting PD control, the feedback state quantity is the attitude angle error and the angular velocity error; combining the first ESO open-loop transfer function to obtain the attitude tracking control model of the pitch channel; combining the influence of the correction network and the extended state observer to obtain the attitude tracking control model;

[0021] The load reduction control mode of the extended state observer specifically includes: adding a normal overload on the basis of the attitude tracking control model feedback; combined with the second ESO open-loop transfer function, the load reduction control model of the pitch channel is obtained; combined with the influence of elastic vibration, the load reduction control model is obtained.

[0022] Optionally, stability derivation is performed on the extended state observer in the attitude tracking control mode to obtain a first ESO open-loop transfer function, specifically including: performing a Laplace transform on the extended state observer; obtaining an open-loop transfer function of the launch vehicle ascent stage dynamics model according to the pitch channel linearization equation, the feedback state quantity and the Laplace transformed extended state observer; according to the open-loop transfer function of the launch vehicle ascent stage dynamics model, the amplitude margin and the phase margin are calculated by drawing a Bode diagram, and stability derivation is performed on the extended state observer in the attitude tracking control mode; according to the load reduction control model of the pitch channel, the first ESO open-loop transfer function is obtained; stability derivation is performed on the extended state observer in the load reduction control mode to obtain a second ESO open-loop transfer function;

[0023] Calculating Normal Overload It is the apparent acceleration along the normal direction of the rocket body measured by the inertial device; combining the attitude tracking control model and the open-loop transfer function of the launch vehicle ascent phase dynamics model, the influence of the wind attack angle on the overload is regarded as interference to obtain the load reduction control model; the open-loop transfer function of the launch vehicle ascent phase dynamics model; according to the open-loop transfer function of the launch vehicle ascent phase dynamics model, the amplitude margin and phase margin are calculated by drawing the Bode diagram, and the stability of the extended state observer under the load reduction control mode is derived.

[0024] Optionally, a first preset parameter adjustment is performed on the extended state observer in the attitude tracking control mode according to the first ESO open-loop transfer function and the feedback control law to obtain a first control result feedback control law part of the launch vehicle ascent phase, specifically including: obtaining a nominal small disturbance coefficient according to a nominal launch vehicle ascent phase dynamics model and trajectory ;make ,in, is the first preset parameter to be adjusted; by adjusting , so that the cutoff frequency of the rocket control system meets the first preset threshold; ,in is the observer bandwidth; first take , get the estimated situation, and then gradually adjust the first preset parameter to reach the second preset threshold; the adjustment range of the first preset parameter is .

[0025] Optionally, a second preset parameter adjustment is performed on the extended state observer in the load reduction control mode according to the second ESO open-loop transfer function and the feedback control law to obtain a second control result of the launch vehicle ascent phase, specifically including: obtaining a nominal small disturbance coefficient according to the nominal launch vehicle ascent phase dynamics model and trajectory ;make ,in, is the second preset parameter to be adjusted; by adjusting , so that the cutoff frequency of the rocket control system meets the third preset threshold; design the parameter transition profile so that the gain transition before and after entering the load reduction control mode is smooth; let ,in is the observer bandwidth; first take , get the estimated situation, and then gradually adjust the second preset parameter to reach the fourth preset threshold; the adjustment range of the second preset parameter is ; Take the overload feedback coefficient Satisfy the closed-loop stability condition; take Make By adjusting , and obtain the second control result of the ascent phase of the carrier rocket.

[0026] A launch vehicle ascent stage control system, comprising:

[0027] A data acquisition module is used to acquire the rigid body dynamics data and elastic vibration data of the ascent stage of the launch vehicle;

[0028] A model building module is connected to the data acquisition module and is used to: build a rigid body dynamics model of the ascent stage of a carrier rocket according to the rigid body dynamics data of the ascent stage; build an elastic vibration model of the ascent stage of a carrier rocket according to the elastic vibration data of the ascent stage; the rigid body dynamics model of the ascent stage of a carrier rocket and the elastic vibration model of the ascent stage of a carrier rocket constitute a dynamics model of the ascent stage of a carrier rocket;

[0029] A linearization processing module, connected to the model building module, is used to rewrite the dynamic model of the launch vehicle's ascending stage into a velocity coordinate system and perform linearization processing to obtain a small disturbance linearization model;

[0030] An extended state observer construction module is connected to the linearization processing module and is used to obtain an extended state observer of the rocket system according to the small disturbance linearization model; the working modes of the extended state observer include: an attitude tracking control mode, and / or a load reduction control mode;

[0031] A stability derivation module is connected to the extended state observer construction module and is used to: perform stability derivation on the extended state observer in the attitude tracking control mode to obtain a first ESO open-loop transfer function; perform stability derivation on the extended state observer in the load reduction control mode to obtain a second ESO open-loop transfer function;

[0032] A parameter adjustment module is connected to the stability derivation module and is used to: perform a first preset parameter adjustment on the extended state observer in the attitude tracking control mode according to the first ESO open-loop transfer function and the feedback control law to obtain a first control result of the launch vehicle ascent phase; the first preset parameter includes: controller frequency and controller damping ratio ; According to the second ESO open-loop transfer function and the feedback control law, the second preset parameters of the extended state observer in the load reduction control mode are adjusted to obtain the second control result of the launch vehicle ascent phase; the second preset parameters include: controller frequency , controller damping ratio and overload feedback factor ; The first control result of the carrier rocket's ascent phase and the second control result of the carrier rocket's ascent phase are used to characterize the control result of the carrier rocket's ascent phase.

[0033] An electronic device comprises a memory, a processor and a computer program stored in the memory and executable on the processor, wherein the processor implements the method for controlling the ascent phase of a carrier rocket when executing the computer program.

[0034] A non-transitory computer-readable storage medium stores a computer program, which implements the launch vehicle ascent stage control method when executed.

[0035] In an embodiment of the present invention, in the active disturbance rejection control, the extended state observer (ESO) estimates the total disturbance in the process by comparing the control quantity applied in the dynamics and its influence on the dynamics, and then compensates the total disturbance torque in the control torque, thereby eliminating the influence of the disturbance on the control, thereby improving the accuracy of attitude stability tracking and load reduction control during the flight of the ascent phase of the launch vehicle. BRIEF DESCRIPTION OF THE DRAWINGS

[0036] Figure 1 A schematic flow chart of a method for controlling the ascent phase of a launch vehicle provided by an embodiment of the present invention;

[0037] Figure 2 A schematic diagram of the principle of multi-target interference estimation based on an extended state observer provided in an embodiment of the present invention;

[0038] Figure 3 A block diagram of attitude tracking control based on auto-disturbance rejection provided in an embodiment of the present invention;

[0039] Figure 4 A block diagram of active load reduction control based on active disturbance rejection provided in an embodiment of the present invention;

[0040] Figure 5 A schematic diagram of the effect of overload feedback on stability provided by an embodiment of the present invention;

[0041] Figure 6 Comparison of posture tracking effects of two control methods provided in embodiments of the present invention;

[0042] Figure 7 A schematic diagram of an overload feedback gain variation curve provided by an embodiment of the present invention;

[0043] Figure 8 A schematic diagram of posture tracking comparison provided by an embodiment of the present invention;

[0044] Fig. 9 A schematic diagram of aerodynamic load comparison provided in an embodiment of the present invention.

[0045] Data acquisition module-100, model building module-200, linear processing module-300, extended state observer building module-400, stability derivation module-500, parameter adjustment module-600. DETAILED DESCRIPTION

[0046] The technical solutions in the embodiments of the present invention will be described clearly and completely below in conjunction with the drawings in the embodiments of the present invention. Obviously, the described embodiments are only part of the embodiments of the present invention, rather than all the embodiments.

[0047] The purpose of the present invention is to provide a control method, system and equipment for the ascent phase of a launch vehicle, so as to solve the problem of low accuracy of attitude stability tracking and load reduction control during the flight of the ascent phase of an existing launch vehicle.

[0048] See also Figure 1 , Step S1: Acquire the rigid body dynamics data and elastic vibration data of the ascent stage of the launch vehicle.

[0049] Step S2: constructing a rigid body dynamics model of the ascent stage of a carrier rocket according to the rigid body dynamics data of the ascent stage; constructing an elastic vibration model of the ascent stage of a carrier rocket according to the elastic vibration data of the ascent stage; specifically including.

[0050] Step S21: . (1)

[0051] Where: is the launch system position vector; is the launch system velocity vector; For quality; is the angular velocity vector, are the three-axis angular velocities respectively; is the three-axis attitude angle; is the thrust vector, is the thrust torque vector; is the aerodynamic force vector, is the aerodynamic moment vector; is the gravitational acceleration vector; is the Coriolis acceleration vector; is the implicate acceleration vector; is the function of the consumption in seconds over time; is the inertia matrix.

[0052] Step S22: constructing an elastic vibration model of the ascent stage of a carrier rocket according to the elastic vibration data of the ascent stage, specifically including: . (2)

[0053] Where: For the pitch channel Generalized coordinates of elastic vibration of order modes; They are pitch channel Natural frequencies and damping coefficients of order modes; They represent the pitch channel engine swing angle and angular acceleration respectively; and They are the pitch channel engine swing angle and angular acceleration Influence coefficient of order mode vibration.

[0054] Step S3: The rigid body dynamics model of the launch vehicle's ascent stage and the elastic vibration model of the launch vehicle's ascent stage constitute a dynamics model of the launch vehicle's ascent stage; specifically including.

[0055] The elastic vibration is coupled with the dynamic model of the launch vehicle's ascent phase by affecting the attitude angle and angular velocity sensed by the inertial group; the deviation of the elastic angle and angular velocity measured at the inertial group is.

[0056] . (3)

[0057] Where: For the pitch channel The modal slope of the order elastic mode at the inertia group.

[0058] The dynamic model of the ascent phase of a launch vehicle includes two parts: the rigid body part and the elastic dynamics. With the launch point of the rocket as the origin, a launch coordinate system is established. The x-axis points to the launch direction in the horizontal plane of the launch point, the y-axis is perpendicular to the horizontal plane of the launch point and points upward, and the z-axis is determined by the right-hand rule. For rockets with a high aspect ratio, the stability of elastic vibration is extremely important. Therefore, when establishing the dynamic model of the ascent phase of the rocket, in addition to the rigid body motion, the elastic vibration of the rocket body is also considered. The generalized coordinates are used to describe the elastic vibration model. The elastic vibration is coupled with the dynamic model of the ascent phase of the launch vehicle by affecting the attitude angle and angular velocity sensitive to the inertial group. The calculation process of the elastic angle and angular velocity deviation measured at the inertial group is. The dynamic model of the ascent phase of the launch vehicle can be constructed into a virtual simulation environment by computer programming using integral methods such as Runge-Kutta, which is used to subsequently verify the effectiveness of the control algorithm. This model is only used to verify the effect of the control algorithm. But it is also a necessary link in the design of the control algorithm, which explains the object processed by the control algorithm, as a simulation dynamic environment and the control algorithm together form a closed-loop system. In order to design the dynamic model of the launch vehicle's ascent phase and analyze the stability of the closed-loop system, the pitch channel is taken as an example, and the above equation is rewritten into the velocity coordinate system and linearized to obtain the formula. The formula is a small disturbance linearization equation, and the controller design and closed-loop stability analysis can be performed based on the equation.

[0059] Step S4: rewriting the dynamic model of the launch vehicle's ascent phase into a velocity coordinate system and performing linearization processing to obtain a small-disturbance linearization model; specifically including: rewriting the dynamic model of the launch vehicle's ascent phase into a velocity coordinate system to perform control algorithm design and closed-loop system stability analysis and perform linearization to obtain the small-disturbance linearization model; specifically including.

[0060] . (4)

[0061] Where: They represent the angle of attack deviation, attitude angle deviation and engine swing angle from the nominal trajectory respectively; are the nominal angle of attack and trajectory inclination; They represent the partial derivative of the normal force coefficient with respect to the angle of attack, the partial derivative of the pitching moment coefficient with respect to the pitching angular velocity, and the partial derivative of the pitching moment coefficient with respect to the angle of attack respectively; Indicates the aircraft reference area; Indicates the reference length of the aircraft; represents the dynamic pressure, and its expression is ; They represent the thrust application point and the distance of the center of mass of the aircraft relative to the head respectively; Indicates interference force; represents the influence of disturbance on posture motion, is the disturbance torque, for The moment of inertia of the shaft.

[0062] Step S5: obtaining an extended state observer of the rocket system according to the small disturbance linearization model; the working mode of the extended state observer includes: an attitude tracking control mode, and / or a load reduction control mode; specifically including: . (5)

[0063] Where: is the small disturbance coefficient of the control quantity under the nominal state; is the observer parameter, by setting the observer parameter reasonably.

[0064] make , (6).

[0065] The attitude tracking control mode of the extended state observer specifically includes: Using PD control, the feedback state quantity is the attitude angle error And the angular velocity error , the calculation formula is as follows. (7). Where: is the rudder deflection generated by PD control. Combined with the first ESO open-loop transfer function, the attitude tracking control model of the pitch channel is obtained, and the calculation formula is as follows. (8). Where: is an adjustable parameter, Preset parameters for the controller; related to rocket dynamics, its physical meaning is the angular acceleration produced by unit rudder deflection under nominal conditions.

[0066] Combining the effects of the correction network and the extended state observer, the attitude tracking control model is obtained, and the calculation formula is as follows. (9). Where: represents the transfer function of the correction network. The load reduction control mode of the extended state observer specifically includes: adding normal overload to the attitude tracking control model The feedback is calculated as follows. (10). Combined with the second ESO open-loop transfer function, the load reduction control model of the pitch channel is obtained, and the calculation formula is as follows. (11). Where: is an adjustable parameter. Combined with the influence of elastic vibration, the load reduction control model is obtained, and the calculation formula is as follows. .

[0067] In one example, the dynamic model of the launch vehicle's ascent phase mainly needs to consider two control tasks (or goals). The first task is attitude angle tracking, that is, responding to the attitude angle command of the guidance loop; the second task is load reduction control during flight in a windy area, which reduces the flight load through control to prevent structural damage. Starting from these two control tasks, the embodiment of the present invention introduces the control system framework design based on the extended state observer.

[0068] The control framework based on the extended state observer is shown in Figure 2. According to formulas (3) and (4), the linearization equation of the pitch channel is obtained.

[0069] . (13)

[0070] Where: is the small disturbance coefficient, corresponding to the coefficient before each linearized variable in formula (4), which respectively represents the influence of various deviations on the center of mass motion and attitude motion. . Will interfere Assuming the expansion state, formula (13) can be rewritten as.

[0071] (14). Where: is the unknown disturbance change rate. Therefore, the extended state observer is designed as (5). By setting the observer parameters reasonably, the above observer can be guaranteed to converge. That is, Thus there is Formula (6) is the design of the extended state observer for the rocket system. At this time, the extended state observer of the rocket system becomes a double-integral series system, and two different feedback control methods can be used for two different control tasks.

[0072] For attitude tracking control mode.

[0073] In addition to the rigid body control task, the elastic vibration is coupled with the control system by affecting the attitude angle and angular velocity that the inertial group is sensitive to. The control system outputs the swing angle command of the elastic frequency point according to the sensitive attitude angle information, which will further aggravate the elastic vibration of the corresponding frequency and form positive feedback. When the swing angle command caused by the elastic vibration reaches saturation, the control system will be unable to correct the attitude error, resulting in flight divergence. Therefore, it is necessary to add a correction network to the control loop to greatly attenuate the sensitivity of the control system at the elastic frequency point. In engineering, notch filters and low-pass filters are usually used to attenuate elastic frequencies of various orders.

[0074] Traditional PD control can directly connect a correction network in series on the main path, but in the design of control systems based on ADRC, the input of the extended state observer contains the control quantity, so it is necessary to consider the impact of the correction network on the observer, and the control equation becomes. After considering the servo model, the control frame Figure 3 as follows.

[0075] The specific control process is to measure the attitude angle error and attitude angular rate error signal during the flight according to the sensor (corresponding to the figure ), the measured attitude angle error and the command attitude angle error (corresponding to the , usually 0), and then combined with the measured attitude angular rate error, the PD control command is obtained using the formula; at the same time, the ESO is based on the control command of the previous moment (i.e., the command rudder deviation, corresponding to the figure ) and the attitude angle error at the current moment, use the formula to estimate the interference, and the estimated result is shown in the formula. Finally, the control command at the current moment is calculated according to the formula .

[0076] For load shedding control mode.

[0077] With the feedback about the normal overload, the normal overload can be reduced to a certain extent. Considering the elastic dynamics model and the servo model, the control frame Figure 4 The specific control process is as follows: According to the attitude angle error and attitude angular rate error signal measured by the sensor during flight (corresponding to the figure ), the measured attitude angle error and the command attitude angle error (corresponding to the , usually 0), combined with the measured attitude angular rate error and the normal overload measured by the inertial combination unit , use the formula to get the PD control command considering overload feedback; at the same time, ESO uses the formula to estimate the interference based on the control command at the previous moment and the attitude angle error at the current moment. The estimated result is shown in the formula. Finally, the control command at the current moment is calculated according to the formula .

[0078] Step S6: Perform stability derivation on the extended state observer in the attitude tracking control mode to obtain a first ESO open-loop transfer function; perform stability derivation on the extended state observer in the load reduction control mode to obtain a second ESO open-loop transfer function; specifically including: Performing Laplace transform on the extended state observer, and the calculation formula is as follows. (15).

[0079] Where: (16). Based on the linearized equation of the pitch channel, the feedback state quantity and the extended state observer after Laplace transformation, the open-loop transfer function of the launch vehicle ascent phase dynamics model is obtained, and the calculation formula is as follows.

[0080] (17). Where: To correct the network transfer function, it consists of five notch filters and a low-pass filter, corresponding to the suppression of the first and second order elastic modes and the modes above the second order; is the transfer function of the servo system, which is equivalent to the second-order model; The rocket body dynamics model.

[0081] According to the open-loop transfer function of the launch vehicle ascent phase dynamics model, the amplitude margin and phase margin are calculated by drawing the Bode diagram, and the stability of the extended state observer in the attitude tracking control mode is derived. According to the load reduction control model of the pitch channel, the first ESO open-loop transfer function is obtained, and the calculation formula is as follows. (18). Where: For arrive transmission; For arrive The transfer function of .

[0082] The stability of the extended state observer under the load reduction control mode is derived, and the second ESO open-loop transfer function is obtained. The calculation formula is as follows.

[0083] (19)Where: . Calculate normal overload is the apparent acceleration along the normal direction of the arrow body measured by the inertial device, and the calculation formula is as follows. . (20) Where: They represent the partial derivatives of normal acceleration with respect to angle of attack and rudder deflection respectively; is the distance between the inertial measurement unit and the center of mass.

[0084] Combining the attitude tracking control model with the open-loop transfer function of the launch vehicle ascent phase dynamics model, the influence of wind attack angle on overload is regarded as interference, and the load reduction control model is obtained. The calculation formula is as follows.

[0085] (21) The open-loop transfer function of the launch vehicle ascent phase dynamics model is calculated as follows. (twenty two).

[0086] According to the open-loop transfer function of the dynamic model of the launch vehicle's ascent phase, the amplitude margin and phase margin are calculated by drawing the Bode diagram, and the stability of the extended state observer in the load reduction control mode is derived.

[0087] In one example, the linearized dynamic model is used to perform stability analysis of the control system. Stability analysis is also one of the necessary steps in controller design. In step S5, the general form of the controller under two different control tasks has been derived. According to the two different control tasks in the ascending section, the stability of the two different controller architectures is derived to obtain two open-loop transfer functions considering ESO.

[0088] In the previous step, it was analyzed that when the observation is accurate, it can be simplified to a series integral type system. However, when performing stability analysis, this assumption cannot be made. The most stringent situation needs to be considered, that is, the observation process cannot be ignored. The stability analysis uses the frequency domain method. Next, the stability of the control system after the closed control loop will be discussed in different situations.

[0089] 1. For the attitude tracking control task, by combining formulas (7), (13), and (16), the relative stability of the system is calculated by analyzing the open-loop transfer function, and the open-loop transfer function of the system (17) is obtained.

[0090] Where: To correct the network transfer function, it consists of five notch filters and a low-pass filter, corresponding to the suppression of the first and second order elastic modes and the modes above the second order; is the transfer function of the servo system, which is equivalent to the second-order model; is the rocket body dynamics model. The relevant expressions are as follows.

[0091] .

[0092] Where: are the natural frequency and damping ratio of the low-pass filter respectively; For the The center frequency of the notch filter is is the notch coefficient, Indicates the attenuation factor at the center frequency; Respectively represent the natural frequency and damping ratio of the servo system. At present, the open-loop transfer function form of the self-disturbance rejection control system has been established. In order to analyze its stability, it is also necessary to adjust the controller parameters. The adjustment parameters are: controller parameters , observer parameters .

[0093] Controller Parameters According to the nominal working conditions Interpolation is determined, where is the interpolation independent variable, which can be time, speed, Mach number or altitude, etc. If the model is more accurate, it can also be calculated according to formula (13) Real-time calculation With the open-loop transfer function, the amplitude margin and phase margin of the system can be calculated by drawing the Bode diagram, and then the relative stability of the system can be evaluated.

[0094] For load shedding control mode.

[0095] Since the relative stability of the attitude angle loop is more concerned during the stability analysis, the open-loop transfer function of the attitude angle loop also needs to be used when solving the open-loop transfer function. Therefore, the transfer function of the overload feedback loop needs to be described by the transfer function related to the attitude angle. Since the attitude angle changes relatively smoothly during the flight of the rocket body, the overload component caused by the accelerated rotation around the center of mass can be ignored during the stability analysis. In the short term, it is assumed that the change in the inertial angle of attack is the same as the change in the attitude angle, that is, .

[0096] At the same time, combined with formulas (11) and (17), the effect of wind angle of attack on overload is regarded as interference, and the calculation process (21) is obtained. When performing stability analysis, it is necessary to consider the influence of each link of the entire system on stability. After considering ESO, feedback control law, rigid body dynamics, servo link, elastic dynamics, and filter network, the system open-loop transfer function is as shown in (22).

[0097] Step S7: According to the first ESO open-loop transfer function and the feedback control law, the first preset parameters of the extended state observer in the attitude tracking control mode are adjusted to obtain the first control result of the launch vehicle ascent phase; the first preset parameters include: controller frequency and controller damping ratio Specific examples include: According to the nominal launch vehicle ascent dynamics model and trajectory, the nominal small disturbance coefficient is obtained .make ,in, is the first preset parameter to be adjusted. , so that the cutoff frequency of the rocket control system meets the first preset threshold. ,in is the observer bandwidth. First, take , get the estimated situation, and then gradually adjust the first preset parameter to reach the second preset threshold; the adjustment range of the first preset parameter is .

[0098] Step S8: According to the second ESO open-loop transfer function and the feedback control law, the second preset parameters of the extended state observer in the load reduction control mode are adjusted to obtain the second control result of the launch vehicle ascent phase; the second preset parameters include: controller frequency , controller damping ratio and overload feedback factor Specific examples include: According to the nominal launch vehicle ascent dynamics model and trajectory, the nominal small disturbance coefficient is obtained .make ,in, is the second preset parameter to be adjusted. , so that the cutoff frequency of the rocket control system meets the third preset threshold. Design the parameter transition profile so that the gain transition before and after entering the load reduction control mode is smooth. ,in is the observer bandwidth. First, take , get the estimated situation, and then gradually adjust the second preset parameter to reach the fourth preset threshold; the adjustment range of the second preset parameter is . Take the overload feedback coefficient Satisfy the closed-loop stability condition. Make By adjusting , and obtain the second control result of the ascent phase of the carrier rocket. The first control result of the ascent phase of the carrier rocket and the second control result of the ascent phase of the carrier rocket are used to characterize the control result of the ascent phase of the carrier rocket.

[0099] In one example, step S6 illustrates the stability analysis process of the algorithm. This step will propose a parameter tuning method. The system provides a tuning method for all parameters to be adjusted in the launch vehicle ascent control method based on the anti-disturbance control algorithm. The parameters to be adjusted in the control system based on anti-disturbance control are composed of two parts. The first part is the parameters of the extended state observer shared by the two control tasks; the second part is the control part parameters, which are also divided into attitude tracking tasks and load reduction control tasks according to different control tasks. The bandwidth method is used to adjust the parameters of the observer. By configuring the bandwidth of the extended state observer, the estimation of the disturbance converges. According to formula (14), the condition for the convergence of the observer is The eigenvalue of is less than 0, and the characteristic equation of the observer is.

[0100] .

[0101] If all the eigenvalues ​​of the above equation are less than zero, then a new characteristic equation is constructed: ,when When , all eigenvalues ​​are guaranteed to be less than 0. (25).

[0102] when and When , the observer can be guaranteed to converge. At this time, the observer parameters to be adjusted are only Next, different parameter setting methods are given according to different control tasks.

[0103] For attitude tracking control mode.

[0104] The control system parameters to be adjusted are , transform it so that , the parameter to be adjusted is the controller frequency and controller damping ratio , adjusting these two parameters can directly reflect the closed-loop response of the control system. For the control loop of the launch vehicle, due to the coupling of elastic vibration and attitude motion, it is necessary to set a filter at the elastic mode frequency point, but this will reduce the phase margin of the system, and the observer bandwidth is also related to the phase margin, and the two jointly affect the margin of the system. Therefore, for the launch vehicle self-disturbance rejection control system considering elastic effects, the selection of the observer bandwidth is limited by the phase margin of the system. Generally speaking , at this time, the control system can be guaranteed to have sufficient margin. The parameter setting method of the launch vehicle active disturbance rejection control system is given as follows.

[0105] Feedback control law part.

[0106] According to the nominal dynamic model and trajectory, the nominal small disturbance coefficient is calculated .make ,in, is the parameter to be adjusted. , so that the cut-off frequency of the control system meets the requirements (usually 1 / 5~1 / 10 of the first-order elastic mode).

[0107] Extended state observer part.

[0108] make ,in is the observer bandwidth. First, take , observe the estimated situation, and then gradually adjust the parameters to achieve the expected performance. During the adjustment process, the stability margin of the system must be ensured. The adjustment range is usually .

[0109] For load shedding control mode.

[0110] The control system parameters to be adjusted are , where the attitude tracking gain The setting method is the same as above, that is, At this time, the parameter to be adjusted is the controller frequency , controller damping ratio And the overload feedback factor Different from the posture tracking task, in the load reduction control task, the influence of overload feedback on stability needs to be considered additionally, and the corresponding overload feedback coefficient also needs to be adjusted additionally.

[0111] It can be seen from formula (20) that the introduction of overload feedback will affect the relative stability of the system. The Bode diagram of the open-loop transfer function before and after overload feedback is as follows: Figure 5 As shown, the horizontal axis is the frequency, the vertical axis is the phase and amplitude of the system, and the punctuation points in the figure represent the amplitude margin and phase margin of the system.

[0112] from Figure 5 It can be seen that after the introduction of overload feedback, the amplitude margin of the low frequency band and the phase margin of the mid-frequency band will be reduced. In order to restore the system margin to the required level, the static error feedback coefficient in the feedback control law can be further reduced. and dynamic feedback error coefficient , improve load shedding control In the proportion of the overall control quantity, the load reduction ratio is increased. The parameter setting rules of the automatic disturbance rejection control algorithm considering overload feedback are given as follows.

[0113] Feedback control law part.

[0114] According to the nominal dynamic model and trajectory, the nominal small disturbance coefficient is calculated .make ,in, is the parameter to be adjusted. , so that the cut-off frequency of the control system meets the requirements (appropriately smaller during load shedding in high wind areas). Design the parameter transition profile so that the gain transition before and after the load shedding control is smooth.

[0115] Extended state observer section.

[0116] make ,in is the observer bandwidth. First, take , observe the estimated situation, and then gradually adjust the parameters to achieve the expected performance. During the adjustment process, the stability margin of the system must be ensured. The adjustment range is usually .

[0117] Overload feedback section.

[0118] Take the overload feedback coefficient Satisfy the closed-loop stability condition. Make By adjusting , so that the load reduction effect meets the design requirements. Due to the use of active disturbance rejection control technology, in the attitude tracking stage, compared with the traditional PID control method, it can eliminate static errors faster and has good transient performance. In a certain wind farm, the comparison between the active disturbance rejection control method (ADRC) and the traditional PID control method is shown in Figure 2. Figure 6 shown.

[0119] The load shedding control works in the high wind area, which is about 6~20km high, and the corresponding flight time is about 20~40 seconds. In this stage, the control loop considers two control methods, namely the traditional proportional differential technology and the anti-disturbance technology based on this paper. Before and after the load shedding control, the overload feedback gain is smoothly transitioned. The parameter transition profile is as follows: Figure 7 As shown; the simulation stop condition is the first-level engine shutdown.

[0120] Figures 8 to 9 The attitude tracking situation and load reduction effect comparison between the method proposed in the embodiment of the present invention (ADRC-LR) and the traditional load reduction control method (PD-LR) in a certain type of wind farm are given.

[0121] In the above simulation results, overall, after the introduction of load reduction control, both control methods can have attitude angle deviations under the feedback of overload, so that the arrow body can balance overload control and attitude tracking; after the load reduction control is completed, under the action of the guidance loop, it gradually returns to the nominal trajectory. Comparing the two methods, ADRC-LR can have a smaller attitude angle deviation and a lower absolute value of the total angle of attack. This is because ADRC has better anti-interference ability and can estimate and compensate for interference during flight through ESO. At the same time, the load reduction ratio using the ADRC-LR algorithm is about 20%, which is slightly higher than the traditional load reduction control method. The above simulation illustrates the effectiveness of the proposed algorithm. Compared with the traditional control system, the control system based on ADRC can have better attitude tracking performance and load reduction performance.

[0122] To sum up, in an embodiment of the present invention, a rigid body dynamics model of the ascent stage of a launch vehicle and an elastic vibration model of the ascent stage of a launch vehicle are constructed; a dynamics model of the ascent stage of a launch vehicle is formed; it is rewritten into a velocity coordinate system and linearized to obtain a small disturbance linearization model; an extended state observer of the rocket system is obtained according to the small disturbance linearization model; stability derivation of the extended state observer is performed to obtain a first ESO open-loop transfer function and a second ESO open-loop transfer function; the first preset parameter is adjusted to obtain a first control result of the ascent stage of the launch vehicle; the second preset parameter is adjusted to obtain a second control result of the ascent stage of the launch vehicle; and in the active disturbance rejection control, the extended state observer (ESO) estimates the total disturbance in this process by comparing the control quantity applied in the dynamics and its influence on the dynamics, and then compensates the total disturbance torque in the control torque, thereby eliminating the influence of the disturbance on the control, thereby improving the accuracy of attitude stability tracking and load reduction control during the flight of the ascent stage of the launch vehicle.

[0123] The various embodiments in this specification are described in a progressive manner, and each embodiment focuses on the differences from other embodiments. The same or similar parts between the various embodiments can be referenced to each other.

[0124] This article uses specific examples to illustrate the principles and implementation methods of the embodiments of the present invention. The description of the above embodiments is only used to help understand the methods and core ideas of the embodiments of the present invention. At the same time, for those skilled in the art, according to the ideas of the embodiments of the present invention, there will be changes in the specific implementation methods and application scope. In summary, the content of this specification should not be understood as limiting the embodiments of the present invention.

Claims

1. A method for controlling the ascent phase of a launch vehicle, characterized in that: include: Acquire rigid body dynamics data and elastic vibration data of the ascent stage of the launch vehicle; construct a rigid body dynamics model of the ascent stage of the launch vehicle according to the rigid body dynamics data of the ascent stage; and construct an elastic vibration model of the ascent stage of the launch vehicle according to the elastic vibration data of the ascent stage; The rigid body dynamics model of the launch vehicle's ascent stage and the elastic vibration model of the launch vehicle's ascent stage constitute a dynamics model of the launch vehicle's ascent stage; the dynamics model of the launch vehicle's ascent stage is rewritten into a velocity coordinate system and linearized to obtain a small disturbance linearization model; Obtaining an extended state observer of the rocket system according to the small disturbance linearization model; The working modes of the extended state observer include: an attitude tracking control mode and / or a load shedding control mode; stability derivation is performed on the extended state observer in the attitude tracking control mode to obtain a first ESO open-loop transfer function; stability derivation is performed on the extended state observer in the load shedding control mode to obtain a second ESO open-loop transfer function; According to the first ESO open-loop transfer function and the feedback control law, the extended state observer in the attitude tracking control mode is adjusted by the first preset parameters to obtain the first control result of the launch vehicle ascent phase; the first preset parameters include: controller frequency and controller damping ratio; According to the second ESO open-loop transfer function and the feedback control law, the second preset parameter adjustment is performed on the extended state observer under the load reduction control mode to obtain the second control result of the ascent phase of the launch vehicle; the second preset parameters include: controller frequency, controller damping ratio and overload feedback coefficient; the first control result of the ascent phase of the launch vehicle and the second control result of the ascent phase of the launch vehicle are used to characterize the control result of the ascent phase of the launch vehicle.

2. The launch vehicle ascending stage control method according to claim 1, characterized in that: The rigid body dynamics model of the launch vehicle's ascent stage is constructed according to the rigid body dynamics data of the ascent stage, specifically including: ； Where: is the launch system position vector; is the launch system velocity vector; For quality; is the angular velocity vector, are the three-axis angular velocities respectively; is the three-axis attitude angle; is the thrust vector, is the thrust torque vector; is the aerodynamic force vector, is the aerodynamic moment vector; is the gravitational acceleration vector; is the Coriolis acceleration vector; is the implicate acceleration vector; is the function of the consumption in seconds over time; is the inertia matrix; The elastic vibration model of the ascent stage of the launch vehicle is constructed according to the elastic vibration data of the ascent stage, specifically comprising: ; Where: For the pitch channel Generalized coordinates of elastic vibration of order modes; They are pitch channel Natural frequencies and damping coefficients of order modes; They represent the pitch channel engine swing angle and angular acceleration respectively; and They are the pitch channel engine swing angle and angular acceleration Influence coefficient of order vibration mode; The rigid body dynamics model of the launch vehicle's ascent stage and the elastic vibration model of the launch vehicle's ascent stage constitute a dynamics model of the launch vehicle's ascent stage, which specifically includes: elastic vibration is coupled with the dynamics model of the launch vehicle's ascent stage by affecting the attitude angle and angular velocity sensed by the inertial group.

3. The launch vehicle ascending stage control method according to claim 2, characterized in that: The dynamic model of the launch vehicle's ascent phase is rewritten into a velocity coordinate system and linearized to obtain a small-disturbance linearized model, specifically including: rewriting the dynamic model of the launch vehicle's ascent phase into a velocity coordinate system to perform control algorithm design and closed-loop system stability analysis and linearization to obtain the small-disturbance linearized model.

4. The launch vehicle ascending stage control method according to claim 3, characterized in that: The extended state observer of the rocket system is obtained according to the small disturbance linearization model, which specifically includes: ； Where: is the small disturbance coefficient of the control quantity under the nominal state; is the observer parameter, by setting the observer parameter reasonably; make , ; The attitude tracking control mode of the extended state observer specifically includes: adopting PD control, the feedback state quantity is attitude angle error and angular velocity error; combining the first ESO open-loop transfer function to obtain the attitude tracking control model of the pitch channel; combining the influence of the correction network and the extended state observer to obtain the attitude tracking control model; The load reduction control mode of the extended state observer specifically includes: adding a normal overload on the basis of the attitude tracking control model feedback; combined with the second ESO open-loop transfer function, the load reduction control model of the pitch channel is obtained; combined with the influence of elastic vibration, the load reduction control model is obtained.

5. The launch vehicle ascending stage control method according to claim 4, characterized in that: The stability of the extended state observer in the attitude tracking control mode is derived to obtain the first ESO open-loop transfer function, which specifically includes: performing Laplace transformation on the extended state observer; obtaining the open-loop transfer function of the launch vehicle ascent stage dynamics model according to the pitch channel linearization equation, the feedback state quantity and the Laplace transformed extended state observer; According to the open-loop transfer function of the launch vehicle ascent stage dynamics model, the amplitude margin and phase margin are calculated by drawing the Bode diagram, and the stability of the extended state observer in the attitude tracking control mode is derived; according to the load reduction control model of the pitch channel, the first ESO open-loop transfer function is obtained; the stability of the extended state observer in the load reduction control mode is derived to obtain the second ESO open-loop transfer function; the normal overload is calculated The apparent acceleration along the normal direction of the rocket body measured by the inertial device; Combining the attitude tracking control model with the open-loop transfer function of the launch vehicle ascent phase dynamics model, the influence of wind attack angle on overload is regarded as interference to obtain the load reduction control model; according to the open-loop transfer function of the launch vehicle ascent phase dynamics model, the amplitude margin and phase margin are calculated by drawing the Bode diagram, and the stability of the expanded state observer under the load reduction control mode is derived.

6. The launch vehicle ascending stage control method according to claim 1, characterized in that: According to the first ESO open-loop transfer function and the feedback control law, the first preset parameter adjustment is performed on the extended state observer in the attitude tracking control mode to obtain the first control result feedback control law part of the launch vehicle ascent phase, which specifically includes: according to the nominal launch vehicle ascent phase dynamics model and trajectory, the nominal small disturbance coefficient is obtained. ;make ,in, is the first preset parameter to be adjusted; by adjusting , so that the cutoff frequency of the rocket control system meets the first preset threshold; ,in is the observer bandwidth; first take , get the estimated situation, and then gradually adjust the first preset parameter to reach the second preset threshold; the adjustment range of the first preset parameter is .

7. The launch vehicle ascending stage control method according to claim 1, characterized in that: According to the second ESO open-loop transfer function and the feedback control law, the second preset parameter adjustment is performed on the extended state observer in the load reduction control mode to obtain the second control result of the launch vehicle ascent phase, specifically including: according to the nominal launch vehicle ascent phase dynamics model and trajectory, the nominal small disturbance coefficient is obtained ;make ,in, is the second preset parameter to be adjusted; by adjusting , so that the cutoff frequency of the rocket control system meets the third preset threshold; design the parameter transition profile so that the gain transition before and after entering the load reduction control mode is smooth; let ,in is the observer bandwidth; first take , get the estimated situation, and then gradually adjust the second preset parameter to reach the fourth preset threshold; the adjustment range of the second preset parameter is ; Take the overload feedback coefficient Satisfy the closed-loop stability condition; take Make By adjusting , and obtain the second control result of the ascent phase of the carrier rocket.

8. A launch vehicle ascent control system, characterized in that: include: A data acquisition module is used to acquire the rigid body dynamics data and elastic vibration data of the ascent stage of the launch vehicle; A model building module is connected to the data acquisition module and is used to: build a rigid body dynamics model of the ascent stage of a carrier rocket according to the rigid body dynamics data of the ascent stage; build an elastic vibration model of the ascent stage of a carrier rocket according to the elastic vibration data of the ascent stage; the rigid body dynamics model of the ascent stage of a carrier rocket and the elastic vibration model of the ascent stage of a carrier rocket constitute a dynamics model of the ascent stage of a carrier rocket; A linearization processing module, connected to the model building module, is used to rewrite the dynamic model of the launch vehicle's ascending stage into a velocity coordinate system and perform linearization processing to obtain a small disturbance linearization model; An extended state observer construction module, connected to the linearization processing module, for obtaining an extended state observer of the rocket system according to the small disturbance linearization model; The working modes of the extended state observer include: posture tracking control mode, and / or load reduction control mode; A stability derivation module is connected to the extended state observer construction module and is used to: perform stability derivation on the extended state observer in the attitude tracking control mode to obtain a first ESO open-loop transfer function; perform stability derivation on the extended state observer in the load reduction control mode to obtain a second ESO open-loop transfer function; A parameter adjustment module is connected to the stability derivation module and is used to: perform a first preset parameter adjustment on the extended state observer in the attitude tracking control mode according to the first ESO open-loop transfer function and the feedback control law to obtain a first control result of the launch vehicle ascent phase; the first preset parameter includes: controller frequency and controller damping ratio ; According to the second ESO open-loop transfer function and the feedback control law, the second preset parameters of the extended state observer in the load reduction control mode are adjusted to obtain the second control result of the launch vehicle ascent phase; the second preset parameters include: controller frequency , controller damping ratio and overload feedback factor ; The first control result of the carrier rocket's ascent phase and the second control result of the carrier rocket's ascent phase are used to characterize the control result of the carrier rocket's ascent phase.

9. An electronic device comprising a memory, a processor, and a computer program stored in the memory and executable on the processor, characterized in that: When the processor executes the computer program, the launch vehicle ascent phase control method as described in claims 1-7 is implemented.

10. A non-transitory computer-readable storage medium having a computer program stored thereon, characterized in that: When the computer program is executed, the launch vehicle ascent phase control method as described in claims 1-7 is implemented.

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