Orbit insertion control method and system for launch vehicle

Abstract

An orbit insertion control method and system for a launch vehicle. For a guidance loop and an attitude loop, equivalent second-order error dynamics models are constructed, respectively; by making full use of the capability of fully characterizing the system design degree of freedom of a second-order fully actuated system theory, comprehensively considering multiple task requirements such as dynamic performance, control constraints, and interference suppression, and introducing an iteration idea into a discrete design framework, hybrid rapid optimization design of guidance loop and attitude loop control gains is achieved, effectively improving overall design efficiency, and design margins can be essentially released, improving the accuracy and performance of orbit insertion control; moreover, guidance loop feedback gain optimization design and attitude command solving are implemented, essentially improving the performance and robustness of orbit insertion control, and attitude loop control design is implemented, essentially ensuring the tracking performance, greatly reducing the conservative property of conventional separate design of dual-loop gains.

Description

A vehicle orbit entry control method and system

[0001] This application claims priority to the Chinese patent application filed with the China Patent Office on October 31, 2023, with application number 202311434117.6 and invention name “A method and system for controlling the entry of a vehicle into orbit”, the entire contents of which are incorporated by reference into this application. Technical Field

[0002] The present invention relates to a carrier orbit insertion control method and system, and in particular to a carrier orbit insertion control method based on parametric design of a second-order all-wheel drive system applicable to the orbit insertion phase of a carrier rocket's final stage or upper stage, belonging to the field of aircraft dynamics and control technology. Background Art

[0003] In order to achieve space missions such as orbit insertion, interplanetary orbit transfer, and orbital maneuvering, launch vehicles represented by the launch vehicle's upper stage, upper stage, and space transport stage are often equipped with one or two high-thrust liquid engines to ensure sufficient active power to reach the required position and speed within a limited time. During the entire flight, the launch vehicle's attitude is adjusted through a servo mechanism or RCS to achieve thrust vector control.

[0004] In the traditional design model for vehicle attitude control systems, guidance design (also known as position control design) and attitude control design are often performed independently after their respective performance targets are assigned. The resulting control commands are then simulated and iterated through the vehicle's six-degree-of-freedom simulation to meet the trajectory accuracy of the vehicle's motion terminal and the servomechanism constraints during flight (such as amplitude constraints and dynamic performance constraints). While this "divide and conquer" attitude control design model facilitates the design of subsystem control algorithms, in principle, the tracking delay of the attitude control subsystem in responding to guidance commands will undoubtedly affect orbital control accuracy. Furthermore, this design model requires pre-assignment of performance targets, resulting in conservative final design specifications.

[0005] Summary of the Invention

[0006] The technical problem solved by the present invention is: to overcome the shortcomings of the existing technology, provide a carrier orbit insertion control method and system, realize the hybrid rapid optimization design of the guidance and attitude dual-loop control gains, effectively improve the overall design efficiency, and can essentially release the design margin and improve the orbit insertion control accuracy and performance.

[0007] The above-mentioned purpose of the present invention is mainly achieved through the following technical solutions:

[0008] A method for controlling a vehicle's orbital insertion comprises the following steps:

[0009] (1) Obtain the launch vehicle's orbital mission parameters, overall parameters, and control design constraints;

[0010] (2) Setting the guidance loop co-morphic parameterization matrix, attitude loop co-morphic parameterization matrix, guidance loop characteristic parameterization matrix and attitude loop characteristic parameterization matrix;

[0011] (3) Calculating the generalized characteristic matrix of the vehicle guidance loop based on the guidance loop co-morphic parameterization matrix and the guidance loop characteristic parameterization matrix; calculating the generalized characteristic matrix of the vehicle attitude loop based on the attitude loop co-morphic parameterization matrix and the attitude loop characteristic parameterization matrix;

[0012] (4) Solve the dual-loop gain design optimization problem based on the vehicle guidance loop generalized characteristic matrix and the vehicle attitude loop generalized characteristic matrix. If the problem has a solution, obtain the optimized guidance loop co-morphic parameterization matrix and attitude loop co-morphic parameterization matrix, as well as the guidance loop characteristic parameterization matrix and attitude loop characteristic parameterization matrix, and proceed to step (5); otherwise, return to step (2);

[0013] (5) Obtain the expected position vector, expected velocity vector and flight state parameters of the carrier at the current moment, and determine whether the carrier meets the orbit entry requirements. If so, the control process ends; otherwise, enter step (6);

[0014] (6) Calculating the control gain matrix of the vehicle guidance loop at the current moment, calculating the desired thrust vector at the current moment based on the control gain matrix, and calculating the desired control thrust and desired yaw-pitch angle vector of the vehicle at the current moment based on the desired thrust vector;

[0015] (7) calculating the roll control torque of the vehicle at the current moment based on the vehicle flight state parameters at the current moment;

[0016] (8) Calculating the yaw-pitch attitude loop control gain matrix of the vehicle at the current moment based on the guidance loop co-morphic parameterization matrix, the attitude loop co-morphic parameterization matrix, and the vehicle flight state parameters at the current moment, and calculating the pitch control torque and yaw control torque of the vehicle at the current moment based on the control gain matrix;

[0017] (9) Control the vehicle into orbit according to the desired control thrust, the vehicle roll control torque, the vehicle pitch control torque, and the yaw control torque of the vehicle, and return to step (5).

[0018] In the above-mentioned vehicle orbit control method, in step (1), the vehicle orbit mission parameters are obtained, including the desired position control accuracy e of the vehicle orbit rss , expected speed control accuracy e vss and control period T;

[0019] Obtain the overall parameters and control design constraints of the vehicle, including the specific impulse I of the vehicle's main engine sp , Maximum thrust of main engine F max , the maximum response speed difference η between attitude and guidance loop, attitude and flexible loop bandwidth multiple constraints and the first-order vibration frequency ω0 of the carrier.

[0020] In the above-mentioned vehicle orbit control method, in step (2), the guidance loop co-state parameterization matrix Z is set r ∈R 3×6 and the attitude loop co-parameterization matrix Z θ ∈R 2×4 ; If it is set for the first time, it is the matrix Z r and Z θ Arbitrary assignment.

[0021] In the above-mentioned vehicle orbit control method, in step (2), the guidance loop characteristic parameterization matrix F is set according to the vehicle overall parameters and control design constraints. r ∈R 6×6 and the attitude loop feature parameterization matrix F θ ∈R 4×4 , and set it to a diagonal structure matrix, that is: F r =diag{s r1 ,s r2 ,...,s r6},F θ =diag{s θ1 ,s θ2 ,...,s θ4},

[0022] Among them, s ri ,s θj , i=1,2,...,6, j=1,2,...,4 are arbitrary real numbers that satisfy the following dual-loop gain design constraints:

[0023] Where η is the maximum response speed difference between attitude and guidance loop; is the attitude and flexible loop bandwidth multiple constraint; ω0 is the first-order vibration frequency of the vehicle.

[0024] In the above-mentioned vehicle orbit insertion control method, in step (3), the generalized characteristic matrix of the vehicle guidance loop is calculated based on the guidance loop co-state parameterization matrix and the guidance loop characteristic parameterization matrix, including:

[0025] Among them, V r is the generalized characteristic matrix of the vehicle guidance loop; Z r F is the guidance loop co-modulation parameterization matrix;r is the characteristic parameterization matrix of the guidance loop.

[0026] In the above-mentioned vehicle orbit insertion control method, step (3) calculates the generalized characteristic matrix of the vehicle attitude loop based on the attitude loop co-state parameterization matrix and the attitude loop characteristic parameterization matrix; including:

[0027] Among them, V θ is the generalized characteristic matrix of the vehicle attitude loop; Z θ is the attitude loop co-morphology parameterization matrix; F θ is the parameterization matrix of the attitude loop feature.

[0028] In the above-mentioned vehicle orbit insertion control method, in step (4), the dual-loop gain design optimization problem is solved based on the generalized characteristic matrix of the vehicle guidance loop and the generalized characteristic matrix of the vehicle attitude loop, satisfying:

[0029] If the problem has a solution, the optimized guidance loop and attitude loop co-parameterization matrix Z is obtained r ,Z θ , guidance loop and attitude loop characteristic parameterization matrix F r ,F θ .

[0030] In the above-mentioned vehicle orbit control method, the vehicle desired position vector at the current moment in step (5) is expressed as r d (t k ), the desired velocity vector is represented by v d (t k ), the flight state parameters of the carrier include mass m(t k ), rolling moment of inertia J x (t k ), yaw moment of inertia J y (t k ), pitch moment of inertia J z (t k ), position vector r(t k ), velocity vector v(t k ), roll angle γ(t k ), roll angular rate Yaw-pitch angle vector θ(t k ) and the yaw-pitch rate vector t k Indicates the current moment.

[0031] In the above-mentioned vehicle orbit control method, the yaw-pitch angle vector θ(t k ) and the yaw-pitch rate vector It is expressed as follows:

[0032] Calculate the current rolling angular velocity ω according to the following formula x (t k ), yaw angular velocity ω y (t k ) and pitch angular velocity ω z (t k ):

[0033] Among them, ψ(t k ) represents the current yaw angle, Indicates the yaw rate at the current moment, Indicates the current pitch angle, Indicates the pitch angle rate at the current moment.

[0034] In the above-mentioned vehicle orbit insertion control method, in step (5), based on the vehicle orbit insertion mission parameters, the vehicle expected position vector, expected velocity vector and vehicle flight state parameters at the current moment, the following formula is used to determine whether the vehicle meets the orbit insertion requirements, that is: k )-r d (t k )||≤e rss ,||v(t k )-v d (t k )||≤e vss

[0035] Among them, e rss is the expected position control accuracy, e vss is the expected speed control accuracy, r d (t k ) is the desired position vector, v d (t k ) is the desired velocity vector, r(t k ) is the position vector, v(t k ) is the velocity vector.

[0036] In the above-mentioned vehicle orbit insertion control method, in step (6), the vehicle guidance loop control gain matrix at the current moment is calculated based on the optimized guidance loop co-state parameterization matrix, the guidance loop characteristic parameterization matrix, the vehicle flight state parameters at the current moment, and the vehicle guidance loop generalized characteristic matrix, including the position gain matrix and the velocity gain matrix, which are expressed as:

[0037] Among them, W r (t k ) is the adjoint matrix of the guidance loop at the current moment, which is calculated by the following formula:

[0038] Where μ is the Earth's gravitational constant, r(t k ) is the vector r(t k ), r(t k ) is the position vector, m(t k ) is mass, K r is the position gain matrix, K v is the velocity gain matrix, V r is the generalized characteristic matrix of the vehicle guidance loop; Z r is the guidance loop co-modulation parameterization matrix; F r is the characteristic parameterization matrix of the guidance loop.

[0039] In the above-mentioned vehicle orbit insertion control method, in step (6), the current moment expected thrust vector is calculated based on the vehicle guidance loop control gain matrix, the current moment expected position vector, the current moment expected velocity vector and the vehicle flight state parameters, including: F d (t k )=F b (t k )+F c (t k )

[0040] Among them, F d (t k ) is the desired thrust vector, F b (t k ),F c (t k ) are the current desired thrust vector compensation term and feedback term, respectively, and are calculated as follows: F c (t k )=K r (t k )(r(t k )-r d (t k ))+K v (t k )(v(t k )-v d (t k ))

[0041] Among them, K r (t k ) is the position gain matrix, K v (t k ) is the velocity gain matrix, μ is the earth's gravitational constant, r d (t k ) is the desired position vector, v d (tk ) is the desired velocity vector, r(t k ) is the position vector, m(t k ) is mass, v(t k ) is the velocity vector.

[0042] In the above-mentioned vehicle orbit insertion control method, step (6) calculates the vehicle's desired control thrust and desired yaw-pitch angle vector at the current moment based on the desired thrust vector at the current moment, including:

[0043] Among them, F d is the desired control thrust, θ d (t k ) is the desired yaw-pitch angle vector, F d,x (t k ),F d,y (t k ),F d,z (t k ) are the desired thrust vector F d (t k ) three-axis components, ψ d (t k ), is the desired yaw angle and the desired pitch angle, calculated according to the following formula:

[0044] In the above-mentioned vehicle orbit insertion control method, step (7) calculates the vehicle roll control torque at the current moment based on the vehicle flight state parameters at the current moment, including:

[0045] Among them, M x (t k ) is the vehicle roll control torque, k ωx is the roll angular velocity gain, α is any positive number that satisfies α∈(0,0.5], is the roll angular rate, J x (t k )Rolling moment of inertia, J y (t k ) is the yaw moment of inertia, J z (t k ) is the pitch moment of inertia, ω x (t k ) is the roll angular velocity, ω y (t k ) is the yaw angular velocity, ω z (t k ) is the pitch angular velocity, t k-1 represents the previous moment, tk Indicates the current moment; ω xc (t k ) is the expected roll angular velocity at the current moment, which is calculated as follows: ω xc (t k )=-k γ γ(t k )-ω y (t k )tanψ(t k )sinγ(t k )-ω z (t k )tanψ(t k )cosγ(t k )

[0046] Among them, γ(t k ) is the roll angle, k γ is the roll angle gain, ψ(t k ) is the yaw angle, is the expected roll angle change rate at the current moment, calculated as follows:

[0047] In the above-mentioned vehicle orbit control method, in step (8), the vehicle yaw-pitch attitude loop control gain matrix at the current moment is calculated based on the optimized attitude loop co-state parameterization matrix and attitude loop characteristic parameterization matrix, as well as the vehicle flight state parameters at the current moment, including the attitude gain matrix K θ (t k )∈R 2×2 And the angular velocity gain matrix K ω (t k )∈R 2×2 , specifically:

[0048] Among them, W θ (t k ) is the current attitude loop adjoint matrix, which is calculated by the following formula:

[0049] Among them, Z θ is the attitude loop co-morphology parameterization matrix; F θ is the parameterized matrix of the attitude loop feature, J(t k ) is the yaw-pitch moment of inertia matrix at the current moment, N(t k ) is the yaw-pitch positive kinematic matrix at the current moment, is the yaw-pitch inverse kinematics derivative matrix at the current moment, which is calculated as follows: J(t k )=diag{J y (t k ),Jz (t k )}

[0050] Among them, γ(t k ) is the roll angle, is the roll angular rate, J y (t k ) is the yaw moment of inertia, J z (t k ) is the pitch moment of inertia, ω x (t k ) is the roll angular velocity, ω y (t k ) is the yaw angular velocity, ω z (t k ) is the pitch angular velocity, ψ(t k ) is the yaw angle, is the yaw rate, t k Indicates the current moment.

[0051] In the above-mentioned vehicle orbit insertion control method, step (8) calculates the pitch control torque and yaw control torque at the current moment based on the vehicle yaw-pitch attitude loop control gain matrix and the vehicle flight state parameters at the current moment, including:

[0052] Among them, M y (t k ) is the pitch control torque, M z (t k ) is the yaw control torque, M b (t k ),M c (t k ) are the current moment control torque vector compensation term and feedback term, respectively, and are calculated by the following formula:

[0053] in, is the expected yaw-pitch angular rate vector and angular acceleration vector at the current moment, n(t k ) is the pitch-yaw antisymmetric compensation vector at the current moment, which is calculated as follows:

[0054] A vehicle orbit insertion control system, comprising:

[0055] Parameter acquisition module, which obtains the vehicle's orbital mission parameters, overall parameters, and control design constraints;

[0056] Matrix design module, setting the guidance loop co-morphic parameterization matrix, attitude loop co-morphic parameterization matrix, guidance loop characteristic parameterization matrix and attitude loop characteristic parameterization matrix;

[0057] a matrix calculation module, which calculates the generalized characteristic matrix of the vehicle guidance loop based on the guidance loop co-morphic parameterization matrix and the guidance loop characteristic parameterization matrix; and calculates the generalized characteristic matrix of the vehicle attitude loop based on the attitude loop co-morphic parameterization matrix and the attitude loop characteristic parameterization matrix;

[0058] an optimization module, which solves a dual-loop gain design optimization problem based on the vehicle guidance loop generalized characteristic matrix and the vehicle attitude loop generalized characteristic matrix. If the problem has a solution, the optimized guidance loop co-morphic parameterization matrix and attitude loop co-morphic parameterization matrix, as well as the guidance loop characteristic parameterization matrix and attitude loop characteristic parameterization matrix are obtained and output to the orbit entry judgment module. If the problem has no solution, the matrix is ​​redesigned by the matrix design module.

[0059] The orbit insertion judgment module obtains the current carrier's expected position vector, expected velocity vector, and carrier flight state parameters, and determines whether the carrier meets the orbit insertion requirements. If so, the control process ends; if not, the command is sent to the expected control thrust calculation module;

[0060] a desired control thrust calculation module, receiving instructions from the orbit insertion determination module, calculating a control gain matrix of the vehicle guidance loop at a current moment, calculating a desired thrust vector at a current moment based on the control gain matrix, and calculating a desired control thrust and a desired yaw-pitch angle vector of the vehicle at a current moment based on the desired thrust vector;

[0061] a roll control torque calculation module, which calculates the roll control torque of the carrier at the current moment according to the carrier flight state parameters at the current moment;

[0062] a pitch and yaw control torque calculation module, which calculates the current vehicle yaw-pitch attitude loop control gain matrix based on the guidance loop co-morphic parameterized matrix, the attitude loop co-morphic parameterized matrix, and the current vehicle flight state parameters, and calculates the current vehicle pitch control torque and yaw control torque based on the control gain matrix;

[0063] The orbit insertion control module performs orbit insertion control of the vehicle according to the desired control thrust of the vehicle, the roll control torque of the vehicle, the pitch control torque of the vehicle, and the yaw control torque of the vehicle.

[0064] Compared with the prior art, the embodiments of the present invention have at least the following beneficial effects:

[0065] (1) In the embodiment of the present invention, equivalent second-order error dynamics models are constructed for the guidance loop and the attitude loop respectively, making full use of the second-order all-wheel drive system theory's ability to fully characterize the system design freedom, comprehensively considering multiple task requirements such as dynamic performance, control constraints, and interference suppression, and introducing the iterative idea into the discrete design framework to achieve a hybrid rapid optimization design of the control gains of the guidance and attitude dual loops, effectively improving the overall design efficiency, and being able to essentially release the design margin and improve the orbit control accuracy and performance.

[0066] (2) In the embodiment of the present invention, a parametric design method for a vehicle orbit insertion guidance loop based on equivalent reference model tracking is adopted. Based on the construction of the trajectory tracking dynamics model, the method fully utilizes the complete characterization of the system design freedom by the second-order all-wheel drive system theory, comprehensively considers the requirements of stable and rapid orbit insertion, attitude tracking delay compensation, interference suppression, etc., realizes the optimization design of the guidance loop feedback gain and the solution of attitude commands, and essentially improves the orbit insertion control performance and robustness.

[0067] (3) In the embodiment of the present invention, a guidance command tracking design method under roll decoupling is adopted. Through the strong stabilization design of the roll channel, its decoupling from the pitch-yaw channel and system order reduction are achieved. At the same time, the attitude tracking performance, the rocket body flexibility suppression and other requirements are comprehensively considered to realize the attitude loop control design, which essentially ensures the tracking performance.

[0068] (4) In the embodiment of the present invention, a dual-loop feedback gain hybrid rapid iteration design method is adopted to decouple the guidance and attitude into a dual second-order loop. The method makes full use of the parametric design of the second-order full-drive system theory to effectively meet the frequency difference design requirements of the attitude tracking loop and the guidance loop, the elastic frequency constraints of the rocket body, etc., effectively improve the overall design efficiency, and at the same time greatly reduce the conservatism of the traditional dual-loop gain separation design. BRIEF DESCRIPTION OF THE DRAWINGS

[0069] FIG1 is a flow chart of a method for controlling a vehicle entering orbit according to an embodiment of the present invention. DETAILED DESCRIPTION

[0070] The present invention will be described in further detail below with reference to the accompanying drawings and specific embodiments:

[0071] As shown in FIG1 , the vehicle orbit insertion control method based on the parametric design of the second-order all-wheel drive system of the present invention includes the following steps:

[0072] (1) Obtain the mission parameters for the launch vehicle into orbit, including the expected position control accuracy e of the launch vehicle into orbit rss and the desired speed control accuracy e vss , control period T.

[0073] (2) Obtain the overall parameters and control design constraints of the vehicle, including the specific impulse I of the vehicle's main engine sp , maximum thrust of main engine F max , the maximum response speed difference η between attitude and guidance loop, the bandwidth multiple constraint of attitude and flexible loop The first-order vibration frequency of the carrier is ω0.

[0074] (3) Set the guidance loop co-state parameterization matrix Z r ∈R 3×6 and the attitude loop co-parameterization matrix Z θ ∈R 2×4 If it is set for the first time, it can be the matrix Z r and Z θ Otherwise, the matrix value obtained by the previous step optimization is used.

[0075] (4) According to the overall parameters of the vehicle and the control design constraints, the guidance loop characteristic parameterization matrix F is set. r ∈R 6×6 and the attitude loop feature parameterization matrix F θ ∈R 4×4 , and set it to a diagonal structure matrix, that is: F r =diag{s r1 ,s r2 ,...,s r6},F θ =diag{s θ1 ,s θ2 ,...,s θ4},

[0076] Among them, s ri ,s θj , i=1,2,...,6, j=1,2,...,4 are arbitrary real numbers that must satisfy the following dual-loop gain design constraints:

[0077] (5) According to the guidance loop co-state parameterization matrix Z r and the guidance loop characteristic parameterization matrix F r , calculate the generalized characteristic matrix V of the vehicle guidance loop according to the following formula: r :

[0078] (6) According to the attitude loop co-state parameterization matrix Z θ and the attitude loop feature parameterization matrix F θ , calculate the generalized characteristic matrix V of the vehicle attitude loop according to the following formula θ :

[0079] (7) According to the generalized characteristic matrix V of the vehicle guidance loop r and the generalized characteristic matrix V of the vehicle attitude loop θ , solve the following dual-loop gain design optimization problem:

[0080] Variation Z r ,F r ,Z θ ,F θ ,make The minimum and satisfies the following formula, namely:

[0081] If the problem has a solution, the optimized guidance loop and attitude loop co-parameterization matrix Z is obtained r ,Z θ , guidance loop and attitude loop characteristic parameterization matrix F r ,F θ , and go to step (8); otherwise, return to step (3).

[0082] (8) Get the current time (t k At time k = 0, 1, 2, ...) the desired position vector r of the vehicle d (t k ) and the desired velocity vector v d (t k ).

[0083] (9) Get the current time (t k At time k = 0, 1, 2, ...) the flight state parameters of the carrier, including the mass m (t k ), rolling moment of inertia J x (t k ), yaw moment of inertia J y (t k ), pitch moment of inertia J z (t k ), position vector r(t k ), velocity vector v(t k ), roll angle γ(t k ), roll angular rate Yaw-pitch angle vector θ(t k ) and the yaw-pitch rate vector And represented by the following formula

[0084] At the same time, the current rolling angular velocity ω is calculated according to the following formula: x (t k ), yaw angular velocity ω y (t k ) and pitch angular velocity ω z (tk ):

[0085] Among them, ψ(t k ) represents the current yaw angle, Indicates the yaw rate at the current moment, Indicates the current pitch angle, Indicates the pitch angle rate at the current moment.

[0086] (10) According to the mission parameters of the carrier into orbit (expected position control accuracy e rss and the desired speed control accuracy e vss ), the current carrier's expected position vector r d (t k ) and the desired velocity vector v d (t k ) and the vehicle flight state parameters (position vector r(t k ), velocity vector v(t k )), the following formula is used to determine whether the carrier meets the requirements for orbital entry: k )-r d (t k )||≤e rss ,||v(t k )-v d (t k )||≤e vss

[0087] If satisfied, the control process ends; otherwise, enters step (11).

[0088] (11) According to the optimized guidance loop co-state parameterization matrix Z r , guidance loop characteristic parameterization matrix F r , the current vehicle flight state parameters and the generalized characteristic matrix V of the vehicle guidance loop r , calculate the current time (t k At time k = 0, 1, 2, ...) the vehicle guidance loop control gain matrix, including the position gain matrix K r (t k )∈R 3×3 and the speed gain matrix K v (t k )∈R 3×3

[0089] Where W r (t k ) is the adjoint matrix of the guidance loop at the current moment, which is calculated by the following formula:

[0090] Where μ is the Earth's gravitational constant, r(t k ),r d (t k ) are vectors r(t k ) and r d (t k ), r(t k ) is the vector r(t k ), r(t k ) is the position vector, .

[0091] (12) According to the current moment vehicle guidance loop control gain matrix (position gain matrix K r (t k ), speed gain matrix K v (t k )), the current moment carrier expected position vector r d (t k ), desired velocity vector v d (t k ) and the flight status parameters of the carrier, calculate the current time (t k At time k = 0, 1, 2, ...) the desired thrust vector F d (t k ): F d (t k )=F b (t k )+F c (t k )

[0092] Where, F b (t k ),F c (t k ) are the current desired thrust vector compensation term and feedback term, respectively, and are calculated as follows: F c (t k )=K r (t k )(r(t k )-r d (t k ))+K v (t k )(v(t k )-v d (t k ))

[0093] In the formula, r(t k ) is the position vector, m(t k ) is mass, v(t k ) is the velocity vector.

[0094] (13) According to the current expected thrust vector F d (t k ), calculate the current time (t k At time k = 0, 1, 2, ...) the vehicle's desired control thrust F d , the desired yaw-pitch angle vector θ d (t k ):

[0095] Where, F d,x (t k ),F d,y (t k ),F d，z (t k The desired thrust vector F d (t k )’s three-axis components, ) are the desired yaw angle and the desired pitch angle, calculated according to the following formula

[0096] (14) According to the current flight state parameters of the carrier, the current time (t k At time k = 0, 1, 2, ...) the vehicle roll channel strong stabilization control torque M x (t k ):

[0097] Where k ωx is the roll angular velocity gain, which can be designed as an arbitrary positive number, and α is any positive number that satisfies α∈(0,0.5]. is the roll angular rate, J x (t k )Rolling moment of inertia, J y (t k ) is the yaw moment of inertia, J z (t k ) is the pitch moment of inertia, ω x (t k ) is the roll angular velocity, ω y (t k ) is the yaw angular velocity, ω z (t k ) is the pitch angular velocity, t k-1 represents the previous moment, t k represents the current moment, ω xc (t k ) is the expected roll angular velocity at the current moment, which is calculated as follows: ωxc (t k )=-k γ γ(t k )-ω y (t k )tanψ(t k )sinγ(t k )-ω z (t k )tanψ(t k )cosγ(t k ).

[0098] Where, γ(t k ) is the roll angle, k γ is the roll angle gain, which can be designed as an arbitrary positive number, ψ(t k ) is the yaw angle, is the expected roll angle change rate at the current moment, calculated as follows:

[0099] In an optional embodiment, when k=0, set ω xc (t k-1 )=0.

[0100] (15) According to the optimized attitude loop co-state parameterization matrix Z θ and the attitude loop feature parameterization matrix F θ , and the current carrier flight state parameters, calculate the current time (t k At time k = 0, 1, 2, ...) the vehicle yaw-pitch attitude loop control gain matrix, including the attitude gain matrix K θ (t k )∈R 2×2 And the angular velocity gain matrix K ω (t k )∈R 2×2 :

[0101] Where W θ (t k ) is the current attitude loop adjoint matrix, which is calculated by the following formula:

[0102] Where, J(t k ) is the yaw-pitch moment of inertia matrix at the current moment, N(t k ) is the yaw-pitch positive kinematic matrix at the current moment, is the yaw-pitch inverse kinematics derivative matrix at the current moment, which is calculated as follows: J(t k )=diag{J y (t k ),Jz (t k )}

[0103] Among them, γ(t k ) is the roll angle, is the roll angular rate, J y (t k ) is the yaw moment of inertia, J z (t k ) is the pitch moment of inertia, ω x (t k ) is the roll angular velocity, ω y (t k ) is the yaw angular velocity, ω z (t k ) is the pitch angular velocity, ψ(t k ) is the yaw angle, is the yaw rate, t k Indicates the current moment.

[0104] (16) According to the current moment vehicle yaw-pitch attitude loop control gain matrix (attitude gain matrix K θ (t k )∈R 2×2 And the angular velocity gain matrix K ω (t k )∈R 2×2 ) and the current carrier flight state parameters, calculate the current time (t k At time k = 0, 1, 2, ...) the pitch and yaw control moments are:

[0105] Where M b (t k ),M c (t k ) are the current moment control torque vector compensation term and feedback term, respectively, and are calculated by the following formula:

[0106] Where, is the expected yaw-pitch angular rate vector and angular acceleration vector at the current moment, n(t k ) is the pitch-yaw antisymmetric compensation vector at the current moment, and is calculated as follows:

[0107] In one optional embodiment, when k=0, set

[0108] (17) Output the current time (t k The desired control thrust F at time k = 0, 1, 2, ...) d and the three-axis control torque M x (t k ),M y (t k ),M z (t k ), perform carrier orbit control, and return to step (8).

[0109] The present invention also provides a vehicle orbit insertion control system, comprising a parameter acquisition module, a matrix design module, a matrix calculation module, an optimization module, an orbit insertion determination module, a desired control thrust calculation module, a roll control torque calculation module, a pitch and yaw control torque calculation module, and an orbit insertion control module, wherein:

[0110] Parameter acquisition module, which obtains the vehicle's orbital mission parameters, overall parameters, and control design constraints;

[0111] Matrix design module, setting guidance loop co-morphic parameterization matrix, attitude loop co-morphic parameterization matrix, guidance loop characteristic parameterization matrix and attitude loop characteristic parameterization matrix;

[0112] a matrix calculation module, which calculates the generalized characteristic matrix of the vehicle guidance loop based on the guidance loop co-morphic parameterization matrix and the guidance loop characteristic parameterization matrix; and calculates the generalized characteristic matrix of the vehicle attitude loop based on the attitude loop co-morphic parameterization matrix and the attitude loop characteristic parameterization matrix;

[0113] an optimization module, which solves a dual-loop gain design optimization problem based on the vehicle guidance loop generalized characteristic matrix and the vehicle attitude loop generalized characteristic matrix. If the problem has a solution, the optimized guidance loop co-morphic parameterization matrix and attitude loop co-morphic parameterization matrix, as well as the guidance loop characteristic parameterization matrix and attitude loop characteristic parameterization matrix are obtained and output to the orbit entry judgment module. If the problem has no solution, the matrix is ​​redesigned by the matrix design module.

[0114] The orbit insertion judgment module obtains the current carrier's expected position vector, expected velocity vector, and carrier flight state parameters, and determines whether the carrier meets the orbit insertion requirements. If so, the control process ends; if not, the command is sent to the expected control thrust calculation module;

[0115] a desired control thrust calculation module, receiving instructions from the orbit insertion determination module, calculating a control gain matrix of the vehicle guidance loop at a current moment, calculating a desired thrust vector at a current moment based on the control gain matrix, and calculating a desired control thrust and a desired yaw-pitch angle vector of the vehicle at a current moment based on the desired thrust vector;

[0116] a roll control torque calculation module, which calculates the roll control torque of the carrier at the current moment according to the carrier flight state parameters at the current moment;

[0117] a pitch and yaw control torque calculation module, which calculates the current vehicle yaw-pitch attitude loop control gain matrix based on the guidance loop co-morphic parameterized matrix, the attitude loop co-morphic parameterized matrix, and the current vehicle flight state parameters, and calculates the current vehicle pitch control torque and yaw control torque based on the control gain matrix;

[0118] The orbit insertion control module performs orbit insertion control of the vehicle according to the desired control thrust of the vehicle, the roll control torque of the vehicle, the pitch control torque of the vehicle, and the yaw control torque of the vehicle.

[0119] An embodiment of the present invention provides a vehicle orbit insertion control method and system. The method first obtains mission parameters, overall vehicle parameters, and control design constraints, then calculates the generalized characteristic matrices of the vehicle guidance loop and attitude loop, then determines whether the dual-loop gain design constraints are met. If not, the generalized characteristic matrices of the guidance loop and attitude loop are corrected. Then, the vehicle's expected position vector, expected velocity vector, and flight state parameters at the current moment are obtained, then determines whether the orbit insertion requirements are met. If so, the control process ends. If not, the vehicle's guidance loop gain matrix, expected thrust vector, expected control thrust, expected yaw-pitch angle vector, roll channel strong stabilization control torque, yaw-pitch attitude loop control gain matrix, and pitch and yaw control torques are calculated at the current moment, and finally, the expected control thrust and three-axis control torque are output. The present invention constructs equivalent second-order error dynamics models for the guidance loop and attitude loop respectively, fully utilizing the second-order all-wheel drive system theory's ability to fully characterize the system's design freedom, comprehensively considering multi-task requirements such as dynamic performance, control constraints, and interference suppression, and introducing iterative thinking into a discrete design framework to achieve a hybrid rapid optimization design of the guidance and attitude dual-loop control gains, effectively improving overall design efficiency, and essentially releasing design margins to enhance orbital control accuracy and performance.

[0120] The above description is only the best specific implementation method of the present invention, but the scope of protection of the present invention is not limited thereto. Any changes or replacements that can be easily thought of by any technician familiar with this technical field within the technical scope disclosed by the present invention should be covered by the scope of protection of the present invention.

[0121] The contents not described in detail in the specification of the present invention belong to the common knowledge of professionals in this field.

Claims

1. A method for controlling a vehicle to enter orbit, characterized in that: The steps include: (1) Obtain the launch vehicle’s orbital mission parameters, overall parameters, and control design constraints; (2) Setting the guidance loop co-morphic parameterization matrix, the attitude loop co-morphic parameterization matrix, the guidance loop characteristic parameterization matrix, and the attitude loop characteristic parameterization matrix; (3) Calculating the generalized characteristic matrix of the vehicle guidance loop according to the guidance loop co-morphic parameterization matrix and the guidance loop characteristic parameterization matrix; calculating the generalized characteristic matrix of the vehicle attitude loop according to the attitude loop co-morphic parameterization matrix and the attitude loop characteristic parameterization matrix; (4) Solving the dual-loop gain design optimization problem according to the vehicle guidance loop generalized characteristic matrix and the vehicle attitude loop generalized characteristic matrix. If the problem has a solution, obtaining the optimized guidance loop co-morphic parameterization matrix and attitude loop co-morphic parameterization matrix, as well as the guidance loop characteristic parameterization matrix and attitude loop characteristic parameterization matrix, and proceeding to step (5); otherwise, returning to step (2); (5) Obtain the expected position vector, expected velocity vector and flight state parameters of the vehicle at the current moment, and determine whether the vehicle meets the orbit entry requirements. If so, the control process ends; otherwise, proceed to step (6); (6) calculating a control gain matrix of the vehicle guidance loop at the current moment, calculating a desired thrust vector at the current moment based on the control gain matrix, and calculating a desired control thrust and a desired yaw-pitch angle vector of the vehicle at the current moment based on the desired thrust vector; (7) calculating the rolling control torque of the vehicle at the current moment according to the vehicle flight state parameters at the current moment; (8) Calculate the yaw-pitch attitude loop control gain matrix of the vehicle at the current moment according to the guidance loop co-morphic parameterization matrix, the attitude loop co-morphic parameterization matrix and the vehicle flight state parameters at the current moment, and calculate the pitch control torque and yaw control torque of the vehicle at the current moment according to the control gain matrix; (9) Control the vehicle into orbit according to the desired control thrust, the vehicle roll control torque, the vehicle pitch control torque and the vehicle yaw control torque of the vehicle, and return to step (5).

2. The method for controlling the vehicle entering orbit according to claim 1, characterized in that: In step (1), the mission parameters for the launch vehicle to orbit are obtained, including the expected position control accuracy e of the launch vehicle to orbit. rss , expected speed control accuracy e vss and control period T; Obtain the overall parameters of the vehicle and the control design constraints, including the specific impulse I of the vehicle's main engine sp , Maximum thrust of main engine F max , the maximum response speed difference η between attitude and guidance loop, the bandwidth multiple constraint of attitude and flexibility loop and the first-order vibration frequency of the carrier ω0.

3. The method for controlling the vehicle entering orbit according to claim 1, characterized in that: In step (2), the guidance loop co-state parameterization matrix Z is set r ∈R 3×6 and the attitude loop co-parameterization matrix Z θ ∈R 2×4 ; If it is set for the first time, it is the matrix Z r and Z θ Any value assigned.

4. The method for controlling the vehicle entering orbit according to claim 1, characterized in that: In step (2), the guidance loop characteristic parameterization matrix F is set according to the vehicle overall parameters and control design constraints. r ∈R 6×6 and the attitude loop feature parameterization matrix F θ ∈R 4×4 , and set it to a diagonal structure matrix, that is: F r =diag{s r1 ,s r2 ,...,s r6 },F θ =diag{s θ1 ,s θ2 ,...,s θ4 }, Among them, s ri ,s θj ,i=1,2,...,6,j=1,2,...,4 are any real numbers that satisfy the following dual-loop gain design constraints: Where η is the maximum response speed difference between attitude and guidance loop; is the attitude and flexible loop bandwidth multiple constraint; ω0 is the first-order vibration frequency of the vehicle.

5. The method for controlling the vehicle entering orbit according to claim 1, characterized in that: In step (3), the generalized characteristic matrix of the vehicle guidance loop is calculated according to the guidance loop co-state parameterization matrix and the guidance loop characteristic parameterization matrix, including: Among them, V r is the generalized characteristic matrix of the vehicle guidance loop; Z r is the guidance loop co-state parameterization matrix; F r Parameterize the matrix for the guidance loop characteristics.

6. The method for controlling the vehicle entering orbit according to claim 1, characterized in that: In step (3), the generalized characteristic matrix of the vehicle attitude loop is calculated according to the attitude loop co-state parameterization matrix and the attitude loop characteristic parameterization matrix; including: Among them, V θ is the generalized characteristic matrix of the vehicle attitude loop; Z θ is the attitude loop co-state parameterization matrix; F θ Parameterize the matrix for the attitude loop feature.

7. The method for controlling the vehicle entering orbit according to claim 1, characterized in that: In step (4), the dual-loop gain design optimization problem is solved based on the generalized characteristic matrix of the vehicle guidance loop and the generalized characteristic matrix of the vehicle attitude loop to satisfy: If the problem has a solution, the optimized guidance loop and attitude loop co-parameterization matrix Z is obtained r ,Z θ , guidance loop and attitude loop characteristic parameterization matrix F r ,F θ .

8. The method for controlling the vehicle entering orbit according to claim 1, characterized in that: The expected position vector of the carrier at the current moment in step (5) is represented by r d (t k ), the desired velocity vector is represented by v d (t k ), the flight state parameters of the carrier include mass m(t k ), rolling moment of inertia J x (t k ), yaw moment of inertia J y (t k ), pitch moment of inertia J z (t k ), position vector r(t k ), velocity vector v(t k ), roll angle γ(t k ), roll angular rate Yaw-pitch angle vector θ(t k ) and the yaw-pitch rate vector t k Indicates the current moment.

9. The method for controlling the vehicle entering orbit according to claim 8, characterized in that: The yaw-pitch angle vector θ(t k ) and the yaw-pitch rate vector It is expressed as follows: Calculate the current rolling angular velocity ω according to the following formula x (t k ), yaw angular velocity ω y (t k ) and pitch angular velocity ω z (t k ): Among them, ψ(t k ) represents the current yaw angle, Indicates the yaw rate at the current moment, Indicates the pitch angle at the current moment, Indicates the pitch angle rate at the current moment.

10. The method for controlling the vehicle entering orbit according to claim 1, characterized in that: In step (5), according to the mission parameters of the vehicle entering orbit, the expected position vector, the expected velocity vector and the flight state parameters of the vehicle at the current moment, it is determined whether the vehicle meets the requirements for entering orbit according to the following formula, that is: k )-r d (t k )||≤e rss ,||v(t k )-v d (t k )||≤e vss Among them, e rss is the expected position control accuracy, e vss is the expected speed control accuracy, r d (t k ) is the desired position vector, v d (t k ) is the desired velocity vector, r(t k ) is the position vector, v(t k ) is the velocity vector.

11. The method for controlling the launch vehicle into orbit according to claim 1, characterized in that: In step (6), the control gain matrix of the vehicle guidance loop at the current moment is calculated according to the optimized guidance loop co-state parameterization matrix, the guidance loop characteristic parameterization matrix, the vehicle flight state parameters at the current moment, and the generalized characteristic matrix of the vehicle guidance loop, including the position gain matrix and the velocity gain matrix, which is expressed as: Among them, W r (t k ) is the adjoint matrix of the guidance loop at the current moment, which is calculated by the following formula: Where μ is the Earth's gravitational constant, r(t k ) is the vector r(t k ), r(t k ) is the position vector, m(t k ) is the mass, K r is the position gain matrix, K v is the velocity gain matrix, V r Guiding the vehicle back Path generalized characteristic matrix; Z r is the guidance loop co-state parameterization matrix; F r Parameterize the matrix for the guidance loop characteristics.

12. The method for controlling the vehicle entering orbit according to claim 1, characterized in that: In step (6), the current moment expected thrust vector is calculated based on the current moment vehicle guidance loop control gain matrix, the current moment vehicle expected position vector, the expected velocity vector and the vehicle flight state parameters, including: d (t k )=F b (t k )+F c (t k ) Among them, F d (t k ) is the desired thrust vector, F b (t k ),F c (t k ) are the current desired thrust vector compensation term and feedback term, respectively, and are calculated by the following formula: F c (t k )=K r (t k )(r(t k )-r d (t k ))+K v (t k )(v(t k )-v d (t k )) Among them, K r (t k ) is the position gain matrix, K v (t k ) is the velocity gain matrix, μ is the earth's gravitational constant, r d (t k ) is the desired position vector, v d (t k ) is the desired velocity vector, r(t k ) is the position vector, m(t k ) is the mass, v(t k ) is the velocity vector.

13. The method for controlling vehicle insertion into orbit according to claim 1, characterized in that: In step (6), the desired control thrust and the desired yaw-pitch angle vector of the vehicle at the current moment are calculated according to the desired thrust vector at the current moment, including: Among them, F d is the desired control thrust, θ d (t k ) is the desired yaw-pitch angle vector, F d,x (t k ),F d,y (t k ),F d,z (t k ) are the desired thrust vector F d (t k )’s three-axis components, is the desired yaw angle and the desired pitch angle, calculated according to the following formula:

14. The method for controlling the vehicle entering orbit according to claim 1, characterized in that: In step (7), the rolling control torque of the carrier at the current moment is calculated according to the flight state parameters of the carrier at the current moment, including: Among them, M x (t k ) is the vehicle rolling control torque, k ωx is the roll angular velocity gain, α is any positive number satisfying α∈(0,0.5], is the roll angular rate, J x (t k )Rolling moment of inertia, J y (t k ) is the yaw moment of inertia, J z (t k ) is the pitch moment of inertia, ω x (t k ) is the rolling angular velocity, ω y (t k ) is the yaw angular velocity, ω z (t k ) is the pitch angular velocity, t k-1 represents the previous moment, t k Indicates the current moment; ω xc (t k ) is the expected roll angular velocity at the current moment, which is calculated as follows: oh xc (t k )=-k γ γ(t k )-oh y (t k )tanψ(t k )sinγ(t k )-oh z (t k )tanψ(t k )cosγ(t k ) Among them, γ(t k ) is the rolling angle, k γ is the roll angle gain, ψ(t k ) is the yaw angle, is the expected roll angle change rate at the current moment, calculated as follows:

15. The method for controlling the vehicle entering orbit according to claim 1, characterized in that: In step (8), the control gain matrix of the yaw-pitch attitude loop of the carrier at the current moment is calculated according to the optimized attitude loop co-state parameterization matrix and the attitude loop characteristic parameterization matrix, as well as the flight state parameters of the carrier at the current moment, including the attitude gain matrix K θ (t k )∈R 2×2 And the angular velocity gain matrix K ω (t k )∈R 2×2 , specifically: Among them, W θ (t k ) is the current attitude loop adjoint matrix, which is calculated by the following formula: Among them, Z θ is the attitude loop co-state parameterization matrix; F θ is the characteristic parameterization matrix of the posture loop, J(t k ) is the yaw-pitch moment of inertia matrix at the current moment, N(t k ) is the yaw-pitch positive kinematics matrix at the current moment, is the yaw-pitch inverse kinematics derivative matrix at the current moment, which is calculated as follows: J(t k )=diag{J y (t k ),J z (t k )} Among them, γ(t k ) is the roll angle, is the roll angular rate, J y (t k ) is the yaw moment of inertia, J z (t k ) is the pitch moment of inertia, ω x (t k ) is the rolling angular velocity, ω y (t k ) is the yaw angular velocity, ω z (t k ) is the pitch angular velocity, ψ(t k ) is the yaw angle, is the yaw rate, t k Indicates the current moment.

16. The method for controlling vehicle orbit insertion according to claim 15, characterized in that: In step (8), the pitch control moment and the yaw control moment at the current moment are calculated according to the vehicle yaw-pitch attitude loop control gain matrix at the current moment and the vehicle flight state parameters at the current moment, including: Among them, M y (t k ) is the pitch control moment, M z (t k ) is the yaw control torque, M b (t k ),M c (t k ) are the control torque vector compensation term and feedback term at the current moment, respectively, and are calculated by the following formula: in, is the expected yaw-pitch angular rate vector and angular acceleration vector at the current moment, n(t k ) is the pitch-yaw antisymmetric compensation vector at the current moment, which is calculated as follows:

17. A vehicle orbit entry control system, characterized in that: include: Parameter acquisition module, which obtains the launch vehicle's orbital mission parameters, overall parameters and control design constraints; A matrix design module is used to set the guidance loop co-morphic parameterization matrix, the attitude loop co-morphic parameterization matrix, the guidance loop characteristic parameterization matrix and the attitude loop characteristic parameterization matrix; A matrix calculation module calculates a generalized characteristic matrix of a vehicle guidance loop according to the guidance loop co-state parameterization matrix and the guidance loop characteristic parameterization matrix; and calculates a generalized characteristic matrix of a vehicle attitude loop according to the attitude loop co-state parameterization matrix and the attitude loop characteristic parameterization matrix; an optimization module, solving a dual-loop gain design optimization problem according to the vehicle guidance loop generalized characteristic matrix and the vehicle attitude loop generalized characteristic matrix, and if the problem has a solution, obtaining the optimized guidance loop co-state parameterization matrix and attitude loop co-state parameterization matrix, as well as the guidance loop characteristic parameterization matrix and attitude loop characteristic parameterization matrix, and outputting them to the orbit entry discrimination module, and if the problem has no solution, redesigning the matrix by the matrix design module; The orbit determination module obtains the expected position vector, expected velocity vector and flight state parameters of the vehicle at the current moment, and determines whether the vehicle meets the orbit requirements. If so, the control process ends; if not, the instruction is sent to the expected control thrust calculation module; an expected control thrust calculation module, receiving the instruction sent by the orbit determination module, calculating the control gain matrix of the vehicle guidance loop at the current moment, calculating the expected thrust vector at the current moment according to the control gain matrix, and calculating the expected control thrust and expected yaw-pitch angle vector of the vehicle at the current moment according to the expected thrust vector; A roll control torque calculation module, which calculates the roll control torque of the carrier at the current moment according to the flight state parameters of the carrier at the current moment; A pitch and yaw control moment calculation module calculates the yaw-pitch attitude loop control gain matrix of the vehicle at the current moment according to the guidance loop co-state parameterization matrix, the attitude loop co-state parameterization matrix and the flight state parameters of the vehicle at the current moment, and calculates the pitch control moment and yaw control moment of the vehicle at the current moment according to the control gain matrix; The orbit insertion control module performs orbit insertion control of the vehicle according to the vehicle desired control thrust, vehicle roll control torque, vehicle pitch control torque and vehicle yaw control torque.

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