A launch vehicle boost phase guidance and control method and apparatus

Abstract

A launch vehicle boost phase guidance and control method and device, comprising the following steps and corresponding modules: (1) obtaining the desired trajectory of the boost flight phase, setting relevant parameters; (2) if the current time reaches the maximum flight time of the boost phase, the control process ends; otherwise, step (3) is entered; (3) obtaining the aerodynamic parameters, vehicle parameters and vibration parameters of the launch vehicle at the current time; (4) calculating the aerodynamic lift, lateral force, drag, elastic vibration disturbance observation vector, engine control thrust, command attitude angle, command angular velocity vector and control moment vector of the launch vehicle at the current time; (5) calculating the pseudo-velocity vector observation value, disturbance angular acceleration observation value, elastic vibration disturbance observation process variable and angular velocity observation vector at the next time; (6) outputting the engine control thrust and control moment vector at the current time for implementing control, and then returning to step (2).

Description

TECHNICAL FIELD

[0001] The present application relates to a kind of launch vehicle boost phase guidance and control method and device, belong to aircraft dynamics and control technical field. BACKGROUND

[0002] The launch vehicle flight process mainly includes boost phase and orbit insertion phase, compared, boost phase mainly occurs in atmosphere, aerodynamic effect is obvious, the influence of the position and attitude motion of launch vehicle cannot be ignored.Therefore, the control purpose of boost flight phase is to reduce the influence of aerodynamic effect and other disturbances, so that the flight trajectory of launch vehicle tracks preset trajectory, provides good initial conditions for orbit insertion phase.Through the thrust vector control mode of main engine and servo mechanism, guidance and control are realized in the whole flight process.

[0003] In the traditional control design mode of launch vehicle boost phase, guidance loop and attitude control loop are usually designed separately according to their respective preset performance indicators, and then six-degree-of-freedom simulation deduction and repeated iteration are carried out to meet the terminal constraints of boost flight and various constraints in flight process (such as servo mechanism constraints, overload constraints, etc.); Further, the guidance loop usually adopts perturbation guidance, and the attitude control loop usually adopts preset program angle tracking control method.This kind of "divide and rule" control method cannot realize the effective matching of guidance control loop and performance indicators, is not conducive to the release of design margin, limits the improvement of overall control performance;Moreover, the control method used is simple and extensive for wind disturbance, structural disturbance and other disturbances in boost flight process, is too conservative, and cannot implement accurate suppression, which essentially limits the improvement of control accuracy. SUMMARY

[0004] The technical problem to be solved by the present application is to overcome the shortcomings of the prior art and improve the tracking accuracy and control performance of the preset trajectory in boost phase.

[0005] The object of the present application is achieved by the following technical solutions:

[0006] A launch vehicle boost phase guidance and control method, comprising:

[0007] (1) obtaining the expected trajectory of boost flight phase, setting control parameters, disturbance acceleration observer parameters, elastic vibration disturbance observer parameters, disturbance angular acceleration observer parameters;

[0008] (2) judging whether the current time reaches the maximum flight time of boost phase, if satisfied, the control process is ended; Otherwise, step (3) is entered;

[0009] (3) obtaining the aerodynamic parameters, body parameters and vibration parameters of launch vehicle at the current time;

[0010] (4) calculating aerodynamic lift, lateral force, drag, elastic vibration disturbance observed vector, engine control thrust, command attitude angle, command angular velocity vector, control moment vector of the current time;

[0011] (5) calculating pseudo-velocity vector observation, disturbance angular acceleration observation, elastic vibration disturbance observation process variable, angular velocity observation vector of the next time;

[0012] (6) outputting engine control thrust and control moment vector of the current time for implementing control, and then returning to step (2).

[0013] Preferably, aerodynamic lift, lateral force, drag of the current time are calculated according to aerodynamic parameters, body parameters of the current time.

[0014] Preferably, elastic vibration disturbance observed vector of the current time is calculated according to elastic vibration disturbance observer parameters, body parameters.

[0015] Preferably, engine control thrust and command attitude angle of the current time are calculated according to drag of the current time, body parameters, aerodynamic parameters, expected trajectory of boost flight stage, control parameters, disturbance acceleration observer parameters, distance from the carrier rocket to the earth.

[0016] Preferably, command angular velocity vector of the current time is calculated according to body parameters of the current time, engine control thrust, aerodynamic parameters, aerodynamic lift, lateral force, control period, and control parameters.

[0017] Preferably, control moment vector of the current time is calculated according to body parameters of the current time, rate of change of command angular velocity vector, aerodynamic moment vector, attitude motion matrix, command attitude angle, disturbance angular acceleration observation, and control parameters.

[0018] Preferably, pseudo-velocity vector observation and disturbance angular acceleration observation of the next time are calculated according to body parameters of the current time, drag, engine control thrust, distance from the carrier rocket to the earth, aerodynamic lift, and control period, disturbance acceleration observer parameters.

[0019] Preferably, elastic vibration disturbance observation process variable and angular velocity observation vector of the next time are calculated according to elastic vibration disturbance observed vector of the current time, body parameters, aerodynamic moment vector, control moment vector, command angular velocity vector, vibration parameters, disturbance angular acceleration observation, elastic vibration disturbance angular acceleration vector, and control period, disturbance angular acceleration observer parameters, elastic vibration disturbance observer parameters.

[0020] A carrier rocket boost stage guidance and control device, comprising:

[0021] The initialization module is used for obtaining a desired trajectory of the boost flight segment, and setting control parameters, disturbance acceleration observer parameters, elastic vibration disturbance observer parameters and disturbance angular acceleration observer parameters.

[0022] The determination module is used for ending the control if the current time reaches the maximum flight time of the boost segment, and otherwise, sequentially calling the parameter acquisition module, the calculation module and the control module in a loop.

[0023] The parameter acquisition module is used for obtaining aerodynamic parameters, body parameters and vibration parameters of the launch vehicle at the current time.

[0024] The calculation module is used for calculating aerodynamic lift, lateral force, resistance, elastic vibration disturbance observation vector, engine control thrust, command attitude angle, command angular velocity vector and control moment vector of the launch vehicle at the current time, and calculating pseudo-velocity vector observation value, disturbance angular acceleration observation value, elastic vibration disturbance observation process variable and angular velocity observation vector at the next time.

[0025] The control module is used for outputting the engine control thrust and the control moment vector at the current time, and implementing the control.

[0026] A computer readable storage medium has computer program instructions stored thereon, which, when loaded and run by a processor, cause the processor to execute the above method.

[0027] Compared with the prior art, the present application has the following beneficial effects:

[0028] (1) The method comprehensively considers the influence of the boost segment aerodynamic effect, fully utilizes the system cascade characteristics, and realizes integrated control instruction design of guidance and control based on backstepping design logic, thereby improving the preset trajectory tracking accuracy and performance.

[0029] (2) The method is based on the composite disturbance observation technology of the extended state observer, and the different characteristics of the wind disturbance, aerodynamic disturbance and other disturbances suffered by the launch vehicle during the boost flight process are effectively observed in a targeted manner, so as to provide input for accurate compensation under the integrated design framework, and effectively reduce the design conservatism.

[0030] (3) The method uses the program angle online solution algorithm, combines the small angle assumption and reasonable simplification, realizes the fast solution of the nonlinear thrust vector equation under the multi-attitude representation characteristics, constructs the analytical mapping relationship between the thrust vector and the command attitude, and ensures the solvability of the integrated design method of guidance and control. BRIEF DESCRIPTION OF DRAWINGS

[0031] Figure 1 The flowchart of the steps of the method is shown in the figure. DETAILED DESCRIPTION

[0032] To make the objects, technical solutions and advantages of the present application clearer, the embodiments of the present application will be further described in detail below with reference to the drawings.

[0033] The present application can be applied to the boost phase trajectory robust tracking control problem of the axisymmetric configuration carrier. The present application will be further described in detail below with reference to the drawings:

[0034] 1) Position dynamics model of the boost phase of the carrier rocket

[0035]

[0036]

[0037] Wherein, r is the position vector of the mass center of the carrier rocket relative to the earth center, η is the pseudo velocity vector, g(η) is the position motion vector, a is the control acceleration vector, a d represents the interference acceleration vector, which is described as follows:

[0038]

[0039]

[0040]

[0041] In the formula, x, y, z are three-axis components of the position vector r respectively, V is the speed of the carrier rocket relative to the ground, r is the distance from the carrier rocket to the earth center, and satisfies σ, θ, γ v are the velocity inclination angle, the track yaw angle and the roll angle of the carrier rocket respectively, m is the mass of the carrier rocket, α, β are the attack angle and the sideslip angle of the carrier rocket respectively, F dx , F dy , F dz are three-axis components of the other interference force F d on the rocket body in addition to the aerodynamic force and the control force, P is the control force, L, N, D represent the aerodynamic lift, the lateral force and the drag on the rocket body respectively, which are described as follows

[0042]

[0043] In the formula, ρ is the atmospheric density, S ref is the aerodynamic reference area, α wp is the additional attack angle caused by the steady wind, α wq is the additional attack angle caused by the shear wind, β wp is the additional sideslip angle caused by the steady wind, β wq is the additional sideslip angle caused by the shear wind, is the aerodynamic lift coefficient, Cf is the side force coefficient. d0 Cf is the side force coefficient.

[0044] 2) Attitude dynamics model of the launch vehicle boost phase

[0045]

[0046]

[0047] where J is the launch vehicle moment of inertia matrix, is the launch vehicle attitude angle vector, ω = [ω x ω y ω z ] T is the launch vehicle attitude angular velocity vector, is the attitude motion matrix, g is the position and attitude coupling motion vector, M a is the aerodynamic moment vector, u, u v , u d are the control angle acceleration vector, elastic vibration interference angle acceleration vector and interference angle acceleration vector, respectively, and are described as follows:

[0048]

[0049]

[0050] u = J -1 M p , u v = J -1 M v , u d = J -1 M d

[0051] In the formula, M p , M v , M d are the control moment vector, elastic vibration interference moment, and interference moment (caused by factors such as tank sloshing and wind interference), respectively, and are described as follows:

[0052]

[0053] 3) Elastic vibration equation of the launch vehicle boost phase

[0054] The first-order elastic vibration of the launch vehicle in the roll, yaw and pitch channels can be described as follows:

[0055]

[0056] where, denotes the three-channel vibration modal generalized coordinates, u j denotes the three-channel vibration input acceleration, ω nj is the three-channel elastic vibration first-order modal circular frequency; ξ j is the three-channel elastic vibration first-order modal damping ratio.

[0057] Further, the three-channel first-order vibration elastic vibration equation can be written as

[0058]

[0059] wherein, denotes the three-channel vibration modal generalized variable, W j is the three-channel elastic coefficient matrix, H is the elastic input matrix, which is described as follows, respectively

[0060]

[0061] At the same time, the three-axis components of the elastic vibration disturbance moment M v can be described as follows:

[0062]

[0063] wherein, is the three-channel vibration disturbance coefficient vector, are the three-channel vibration modal proportional coefficient and vibration modal velocity coefficient, respectively.

[0064] 5) A carrier pose integrated control method considering the dynamic characteristics of servo mechanism includes the following steps, as shown in Figure 1 :

[0065] (1) Obtain the task parameters, including the expected trajectory of the boost flight segment, including the expected flight speed V d (t) and its rate of change the expected speed inclination trajectory σ d (t) and its rate of change the expected track yaw angle trajectory θ d (t) and its rate of change the maximum flight time t max of the boost segment, and satisfies 0≤t≤t max , control period T.

[0066] (2) Set the control parameters, including the position gain matrix the attitude gain matrix the control gain matrix

[0067] (3) Set the disturbance acceleration observer parameters, including the disturbance acceleration observation initial value Pseudo-velocity observation vector initial value Disturbance acceleration input observation matrix and disturbance acceleration output observation matrix Satisfies K a1 ,K a2 > 0.

[0068] (4) Set the elastic vibration disturbance observer parameters, including the roll, yaw and pitch three channels elastic vibration disturbance observation initial value Elastic vibration disturbance observation process variable initial value τ j (t0) = 0, elastic vibration disturbance observation gain vector And elastic vibration disturbance observation gain coefficient

[0069] (5) Set the disturbance angular acceleration observer parameters, including the disturbance angular acceleration observation initial value Angular velocity observation vector initial value Disturbance angular acceleration input observation matrix Satisfies K ω > 0.

[0070] (6) Determine whether the current time reaches the maximum flight time of the boost phase, that is, determine t k = t max , if satisfied, the control process is ended; otherwise, go to step (7).

[0071] (7) Obtain the current time (t k ), k = 0, 1, 2,...), the aerodynamic related parameters of the launch vehicle, including the atmospheric density ρ(t k ), the additional angle of attack α wp (t k ) caused by the steady wind, the additional angle of attack α wq (t k ) caused by the shear wind, the additional side slip angle β wp (t k ) caused by the steady wind, the additional side slip angle β wq (t k ) caused by the shear wind, the aerodynamic lift coefficient Lateral force coefficient Drag coefficient C d0 (t k ), dynamic pressure q(t k ) and characteristic reference area S ref (t k ).

[0072] (8) Obtain the current time (t k ), k = 0, 1, 2,...), the launch vehicle body related parameters, including the mass m(tk ), moment of inertia J(t) k ), position vector r(t) k ), flight speed V(t) k ), velocity tilt angle σ(t) k and rate of change Angle of attack α(t) k ), sideslip angle β(t) k ), track yaw angle θ(t) k and rate of change Tilt angle γ v (t k ), pseudo velocity vector η(t) k ), attitude angle vector Attitude angular velocity vector ω(t) k (), Arrow body center position x cp (t k ) and the position of the center of mass x cm (t k ).

[0073] (9) Obtain the current time (t) k At time k = 0, 1, 2, ..., the launch vehicle vibration-related parameters include the generalized coordinates q of the three vibration modes: roll, yaw, and pitch. j (t k Vibration input acceleration u j (t k The first-order modal angular frequency ω of elastic vibration nj (t k The damping ratio of the first mode of elastic vibration, ξ j (t k Vibration disturbance coefficient vector

[0074] (10) Calculate the current time (t) according to the following formula. k At time k = 0, 1, 2, ..., the aerodynamic lift force L(t) on the launch vehicle k Lateral force N(t) k ) and resistance D(t) k ).

[0075]

[0076] (11) Calculate the current time (t) according to the following formula. k At time k = 0, 1, 2, ..., elastic vibration disturbance observation vector

[0077]

[0078] In the formula, ω x (tk ),ω y (t k ),ω z (t k ) is the attitude angular velocity vector ω(t) k The three-channel components of ).

[0079] (12) Calculate the current time (t) according to the following formula. k At time k = 0, 1, 2, ..., the engine control thrust P(t) k and command attitude angle

[0080] P(t k )=u Px (t k )m(t k )+D(t k )

[0081]

[0082]

[0083] γ vc (t k ) = 0

[0084] In the formula,

[0085]

[0086] In the formula, μ is the Earth's gravitational constant, and r(t) k ) for t k The distance between the launch vehicle's center and the Earth's center at any given time corresponds to the position vector r(t). k The norm of ).

[0087] (13) Calculate the current time (t) according to the following formula. k The command angular velocity vector ω at time k = 0, 1, 2, ... c (t k ).

[0088]

[0089] in, For t k Motion Matrix at Any Time The inverse matrix, and Obtained from the following formula

[0090]

[0091] At the same time, g(t) k ) for t kThe time-coupled pose motion vector is obtained by the following formula

[0092]

[0093] And R(t k ) is obtained by the following formula

[0094]

[0095] In the formula,

[0096]

[0097]

[0098]

[0099]

[0100]

[0101]

[0102]

[0103] Meanwhile, is the command angle vector rate at time t k , which can be obtained by the following formula:

[0104]

[0105] Wherein, when k=0, set

[0106] (14) The control torque vector M k at the current time (t p , k=0, 1, 2,...) is calculated according to the following formula: k

[0107]

[0108] Wherein, is the attitude motion matrix at time t k , M a (t k ) is the aerodynamic moment vector at time t k , which is obtained by the following formula:

[0109]

[0110] is the command angular velocity vector rate at time t k , which can be obtained by the following formula:​

[0111]

[0112] where k = 0, ω c (t k-1 ) = 0.

[0113] (15) The pseudo velocity vector observation value at the next time (t k+1 ), k = 0, 1, 2,...), is calculated according to the following equation and the disturbance angular acceleration observation value

[0114]

[0115] where the control acceleration vector a(t k ) = [a x (t k )a y (t k )a z (t k )] T is obtained from the following equation

[0116]

[0117]

[0118]

[0119] (16) The elastic vibration disturbance observation process variable τ(t k+1 ) and the angular velocity observation vector at the next time (t k+1 ), k = 0, 1, 2,...), are updated according to the following equations

[0120]

[0121]

[0122] where

[0123]

[0124]

[0125]

[0126] where J -1 (t k ) is the inverse matrix of the moment of inertia J(t k ).

[0127] (17) outputting the engine control thrust P(t k ) and the control moment vector M k (t p ) at the current time (t k , k = 0, 1, 2,...), implementing control, and returning to step (6).

[0128] A launch vehicle boost phase guidance and control device comprises:

[0129] An initialization module is configured to obtain a desired trajectory of the boost flight phase, and set control parameters, disturbance acceleration observer parameters, elastic vibration disturbance observer parameters, and disturbance angular acceleration observer parameters.

[0130] A determination module is configured to end the control if the current time reaches a maximum flight time of the boost phase, or to sequentially and cyclically call the parameter obtaining module, the calculation module, and the control module.

[0131] The parameter obtaining module is configured to obtain aerodynamic parameters, rocket body parameters, and vibration parameters of the launch vehicle at the current time.

[0132] The calculation module is configured to calculate aerodynamic lift, lateral force, resistance, elastic vibration disturbance observation vectors, engine control thrust, command attitude angles, command angular velocity vectors, and control moment vectors of the launch vehicle at the current time, and to calculate pseudo-velocity vector observation values, disturbance angular acceleration observation values, elastic vibration disturbance observation process variables, and angular velocity observation vectors at the next time.

[0133] The control module is configured to output the engine control thrust and the control moment vector at the current time, and to implement control.

[0134] A computer readable storage medium has computer program instructions stored thereon, and the computer program instructions, when loaded and run by a processor, cause the processor to execute the above method.

[0135] The contents not described in detail in the specification of the present application are known to those skilled in the art.

[0136] Although the present application has been disclosed with reference to the preferred embodiments above, it is not intended to limit the present application, and any person skilled in the art can make possible changes and modifications to the technical solutions of the present application by using the disclosed methods and technical contents without departing from the spirit and scope of the present application. Therefore, any simple modification, equivalent change and modification made to the above embodiments according to the technical essence of the present application, which does not depart from the technical solutions of the present application, belongs to the protection scope of the technical solutions of the present application.

Claims

1. A guidance and control method for the boost phase of a launch vehicle, characterized in that, include: (1) Obtain the desired trajectory of the boost flight phase and set the control parameters, disturbance acceleration observer parameters, elastic vibration disturbance observer parameters, and disturbance angular acceleration observer parameters; (2) Determine whether the current time has reached the maximum flight time of the boost phase. If it has, the control process ends; otherwise, proceed to step (3). (3) Obtain the current aerodynamic parameters, rocket body parameters, and vibration parameters of the launch vehicle; (4) Calculate the aerodynamic lift, lateral force, drag, elastic vibration disturbance observation vector, engine control thrust, commanded attitude angle, commanded angular velocity vector, and control torque vector of the launch vehicle at the current moment; Elastic vibration disturbance observation vector In the formula, ω x (t k ),ω y (t k ),ω z (t k ) is the attitude angular velocity vector ω(t) k The three-channel components; Engine control thrust P(t) k and command attitude angle They are respectively: P(t k )=u Px (t k )m(t k )+D(t k ) c vc (t k )=0 In the formula, In the formula, μ is the Earth's gravitational constant, and r(t) k ) for t k The distance between the launch vehicle's center and the Earth's center at any given time corresponds to the position vector r(t). k The norm of ). Command angular velocity vector ω c (t k )for: in, For t k Motion Matrix at Any Time The inverse matrix, and Obtained from the following formula g(t k ) for t k The pose-coupled motion vector at any given time is obtained from the following equation. And R(t) k The following formula is used to obtain the result. In the formula, at the same time, For t k The rate of change of the command angle vector at any given time can be obtained by the following formula: Where, when k=0, set (5) Calculate the pseudo velocity vector observation value, disturbance angular acceleration observation value, elastic vibration disturbance observation process variable, and angular velocity observation vector at the next moment; pseudo-velocity vector observations and interference angular acceleration observations They are respectively: Among them, the control acceleration vector a(t) k )=[a x (t k )a y (t k )a z (t k )] T It is obtained from the following formula Elastic vibration disturbance observation process variable τ(t) k+1 ) and angular velocity observation vector They are respectively: In the formula In the formula, J -1 (t k ) is the moment of inertia J(t) k The inverse matrix of ). (6) Output the engine control thrust and control torque vector at the current moment for implementation of control, and then return to step (2).

2. The guidance and control method according to claim 1, characterized in that, Based on the current aerodynamic parameters and rocket body parameters, calculate the aerodynamic lift, lateral force, and drag experienced by the launch vehicle at the current moment.

3. The guidance and control method according to claim 1, characterized in that, Calculate the elastic vibration disturbance observation vector at the current moment based on the elastic vibration disturbance observer parameters and the rocket body parameters.

4. The guidance and control method according to claim 1, characterized in that, Based on the drag, rocket body parameters, aerodynamic parameters, desired trajectory, control parameters, interference acceleration observer parameters, and launch vehicle distance from the Earth's center during the boost phase, calculate the engine control thrust and commanded attitude angle at the current moment.

5. The guidance and control method according to claim 1, characterized in that, Calculate the commanded angular velocity vector at the current moment based on the rocket body parameters, engine control thrust, aerodynamic parameters, aerodynamic lift, lateral force, control period, and control parameters.

6. The guidance and control method according to claim 1, characterized in that, Calculate the control torque vector at the current moment based on the rocket body parameters, the rate of change of the commanded angular velocity vector, the aerodynamic torque vector, the attitude motion matrix, the commanded attitude angle, the observed values ​​of the disturbance angular acceleration, and the control parameters.

7. The guidance and control method according to claim 1, characterized in that, Based on the current parameters of the rocket body, drag, engine control thrust, launch vehicle distance from the Earth's center, aerodynamic lift, control period, and interference acceleration observer parameters, calculate the pseudo velocity vector observation value and interference angular acceleration observation value for the next moment.

8. The guidance and control method according to claim 1, characterized in that, Based on the current elastic vibration disturbance observation vector, rocket body parameters, aerodynamic torque vector, control torque vector, command angular velocity vector, vibration parameters, disturbance angular acceleration observation value, elastic vibration disturbance angular acceleration vector, as well as the control period, disturbance angular acceleration observer parameters, and elastic vibration disturbance observer parameters, calculate the elastic vibration disturbance observation process variables and angular velocity observation vector at the next moment.

9. A computer-readable storage medium having stored thereon computer program instructions, which, when loaded and run by a processor, cause the processor to perform the method as described in any one of claims 1 to 8.

Images