A Method for Spacecraft Attitude Control and RCS Optimal Allocation

Abstract

The present invention discloses a method for spacecraft attitude control and RCS optimal allocation, which relates to the technical field of spacecraft control. Spacecraft parameters and spacecraft attitude control torque commands are obtained; a spacecraft attitude dynamics model is constructed; a state space equation is constructed according to the spacecraft attitude dynamics model; an estimated value of the external disturbance torque acting on the spacecraft is obtained according to the state space equation; an error conversion state space equation is constructed according to the finite-time performance function and the estimated value of the external disturbance torque acting on the spacecraft; a super-twisting sliding mode controller is constructed according to the error conversion state space equation; a spacecraft attitude dynamics model after chattering suppression is constructed according to the super-twisting sliding mode controller and the saturation function; an RCS optimal allocation result is obtained according to the spacecraft attitude control torque command and the spacecraft attitude dynamics model after chattering suppression. The present invention realizes the precise tracking control of the spacecraft attitude within a finite time, and optimizes the control allocation with the minimum command error and the overall optimal energy consumption.

Description

Technical Field

[0001] The present invention relates to the technical field of spacecraft control, and particularly to a method for spacecraft attitude control and RCS optimal allocation. Background Technique

[0002] With the rapid development of science and technology, people's demand for exploring the outer space field is increasing continuously, and more and more research requires precise control of spacecraft. As one of the most important systems of spacecraft, the attitude control subsystem needs to provide stable and precise attitude control for the spacecraft so that the spacecraft can complete various space missions and promote the development of the national space industry.

[0003] Theoretically, the convergence time of the traditional spacecraft attitude control method is infinitely far, and the final attitude tracking error range cannot be determined, which is obviously not applicable to the actual engineering applications and will bring difficulties and challenges to the on-orbit operation of the spacecraft. At the same time, in the traditional attitude control method, in order to cope with large disturbances, relatively large controller gains are often set, which will undoubtedly exacerbate the influence of the chattering problem. In most of the studies on spacecraft attitude control, the problem of control command allocation is not considered, and the attitude control of many spacecrafts requires many groups of reaction control systems (RCS) to achieve. How to achieve the optimal control allocation of RCS is also a problem faced in current engineering practice. Summary of the Invention

[0004] The purpose of the embodiments of the present invention is to provide a method for spacecraft attitude control and RCS optimal allocation, which realizes the precise tracking control of the spacecraft attitude within a finite time and optimizes the control allocation with the minimum command error and the optimal comprehensive energy consumption.

[0005] To achieve the above purpose, the embodiments of the present invention provide the following solutions:

[0006] A method for spacecraft attitude control and RCS optimal allocation, the method for spacecraft attitude control and RCS optimal allocation operates based on the predetermined performance of the spacecraft, and includes:

[0007] Obtain spacecraft parameters and spacecraft attitude control torque commands; the spacecraft parameters at least include the inertia matrix;

[0008] Construct a spacecraft attitude dynamics model according to the spacecraft parameters;

[0009] Construct a state space equation according to the spacecraft attitude dynamics model; design a disturbance observer according to the state space equation, and observe and obtain the estimated value of the external disturbance torque received by the spacecraft;

[0010] Construct a finite-time performance function according to the required control performance index; construct an error transformation state-space equation according to the finite-time performance function and the estimated value of the external disturbance torque acting on the spacecraft;

[0011] Construct a super-twisting sliding mode controller according to the error transformation state-space equation;

[0012] Construct a saturation function according to the sliding mode surface; construct a spacecraft attitude dynamics model after suppressing chattering according to the super-twisting sliding mode controller and the saturation function;

[0013] Obtain the RCS optimal allocation result according to the spacecraft attitude control torque command and the spacecraft attitude dynamics model after suppressing chattering; the RCS optimal allocation result includes the on-off state of any one RCS thruster.

[0014] Optionally, construct a spacecraft attitude dynamics model according to the spacecraft parameters, specifically including:

[0015]

[0016]

[0017] Among them, represents the roll in the inertial frame, ψ represents the pitch in the inertial frame, γ represents the yaw angle in the inertial frame, θ ∈ R 3×1 represents the spacecraft attitude angle, represents the first derivative of θ, represents the first derivative of ω, ω ∈ R 3×1 represents the spacecraft attitude angular velocity, M is the three-axis torque acting on the spacecraft;

[0018] Construct a state-space equation according to the spacecraft attitude dynamics model, specifically including:

[0019]

[0020]

[0021] A = -J -1 ω × J;

[0022] B = J -1 ;

[0023] Among them, x1 = θ, x2 = ω, A and B represent the equality sign, represents the first derivative of x2, represents the first derivative of x2, u ∈ R 3×1 represents the control torque, D ∈ R 3×1Represents the estimated value of the external disturbance torque acting on the spacecraft.

[0024] Optionally, a disturbance observer is designed according to the state - space equation to obtain the estimated value of the external disturbance torque acting on the spacecraft, specifically including:

[0025]

[0026] where, L A =diag(L A1 L A2 L A3 )>0; k Ai =diag(k Ai1 k Ai2 k Ai3 ), i = 1,..., 4; κ Ai , α A are all estimated parameters; e x2 =x2 - x 2d ; x 1d represents the command value of the spacecraft attitude control torque command x1; x 2d represents the command value of the spacecraft attitude control torque command x2; the function sig r (x)=|x| r ·sgn(x);

[0027] The parameter σ=diag(σ1 σ2 σ3) is defined as follows:

[0028]

[0029] where, T Aj is the positive switching time parameter.

[0030] Optionally, a finite - time performance function is constructed according to the required control performance index, specifically including:

[0031] ρ fV (t)=a V3 t 4 +a V2 t 3 +a V1 t 2 +c Vρr t + c Vρ0 ;

[0032] where, a V3 、a V2 and a V1 The expressions of are as follows:

[0033]

[0034] where, the parameter cVρ0 , c Vρr , ρ fV∞ and T fV are all predicted parameters, where t represents time; c Vρ0 is a positive number representing the initial error bound, c Vρr represents the initial change direction of the performance function, ρ fV∞ characterizes the function ρ fV (t) at its steady-state convergence value, T fV represents ρ fV (t) converging to the steady-state value ρ fV∞ of the setting time.

[0035] Optionally, according to the finite-time performance function and the estimated value of the external disturbance torque acting on the spacecraft, an error conversion state-space equation is constructed, specifically including:

[0036]

[0037]

[0038]

[0039] where ξ represents the conversion error, represents the first derivative of ξ, represents of the first derivative, Q and P represent equality signs, is the upper bound coefficient of the finite-time performance function error, and δ is the lower bound coefficient of the finite-time performance function error;

[0040] According to the error conversion state-space equation, a super-twisting sliding mode controller is constructed, specifically including:

[0041]

[0042]

[0043] where x'1 = ξ, represents the first derivative of x′1, represents the first derivative of x′2, represents the diagonal matrix composed of the vector Q, represents x 1d of the second derivative.

[0044] Optionally,

[0045] the sliding mode surface s in the super-twisting sliding mode controller is:

[0046] s = Cx'1 + x'2;

[0047] Hyperbolic approach law is as follows:

[0048]

[0049]

[0050]

[0051] wherein, is the first derivative of s, C is a predicted parameter, x′1, x′2, and v are equation symbols; K1, K2 are controller gains, is the first derivative of v; u d represents the control output command, represents the total disturbance observed by the disturbance observer;

[0052] According to the sliding mode surface, a saturation function is constructed, specifically including:

[0053]

[0054] wherein, δ is a preset parameter.

[0055] Optionally,

[0056] According to the spacecraft attitude control torque command and the spacecraft attitude dynamics model after chattering suppression, an RCS optimization allocation result is obtained, specifically including:

[0057]

[0058] M RCS = F I R;

[0059] wherein, M RCS represents the torque provided by the RCS, M d represents the spacecraft attitude control torque command, α represents the weight coefficient, Q1 ∈ R 3×3 and Q2 ∈ R m×m are preset positive definite matrices; F I represents the coefficient matrix of the torque provided by the RCS, R = [r1, r2,..., r m T represents the on-off states of m RCS thrusters, r i can only take 0 or 1 。

[0060] ​In the embodiments of the present invention, a finite-time performance function is constructed, and an error conversion state-space equation is constructed based on the finite-time performance function and the estimated value of the external disturbance torque received by the spacecraft. The control method with predetermined performance can convert the original tracking error into an unconstrained error, thereby ultimately meeting some set control performance indicators, and can achieve precise tracking control of the spacecraft attitude within a finite time. According to the error conversion state-space equation, a super-twisting sliding mode controller is constructed. The super-twisting sliding mode control can effectively cope with large disturbances and will not cause serious chattering phenomena, and can simultaneously achieve the minimum command error and the comprehensive optimization of energy consumption control allocation. BRIEF DESCRIPTION OF THE DRAWINGS

[0061] In order to more clearly illustrate the technical solutions in the embodiments of the present invention or the prior art, the following will briefly introduce the drawings required in the embodiments. Obviously, the drawings in the following description are only some embodiments of the present invention. For those of ordinary skill in the art, without creative efforts, other drawings can be obtained based on these drawings.

[0062] Figure 1 It is a schematic flow chart of the spacecraft attitude control and RCS optimal allocation method provided by the embodiments of the present invention;

[0063] Figure 2 It is a schematic implementation flow chart of the spacecraft attitude control and RCS optimal allocation method based on predetermined performance provided by the embodiments of the present invention;

[0064] Figure 3 It is a roll angle tracking error curve graph provided by the embodiments of the present invention without considering control allocation;

[0065] Figure 4 It is a pitch angle tracking error curve graph provided by the embodiments of the present invention without considering control allocation;

[0066] Figure 5 It is a yaw angle tracking error curve graph provided by the embodiments of the present invention without considering control allocation;

[0067] Figure 6 It is a spacecraft X-axis torque curve graph provided by the embodiments of the present invention without considering control allocation;

[0068] Figure 7 It is a spacecraft Y-axis torque curve graph provided by the embodiments of the present invention without considering control allocation;

[0069] Figure 8 It is a spacecraft Z-axis torque curve graph provided by the embodiments of the present invention without considering control allocation;

[0070] Figure 9Roll angle tracking error curve considering control allocation provided by an embodiment of the present invention;

[0071] Figure 10 Pitch angle tracking error curve considering control allocation provided by an embodiment of the present invention;

[0072] Figure 11 Yaw angle tracking error curve considering control allocation provided by an embodiment of the present invention;

[0073] Figure 12 Spacecraft X-axis torque curve considering control allocation provided by an embodiment of the present invention;

[0074] Figure 13 Spacecraft Y-axis torque curve considering control allocation provided by an embodiment of the present invention;

[0075] Figure 14 Spacecraft Z-axis torque curve considering control allocation provided by an embodiment of the present invention. Detailed implementation manners

[0076] Next, the technical solutions in the embodiments of the present invention will be clearly and completely described in conjunction with the accompanying drawings in the embodiments of the present invention. Obviously, the described embodiments are only a part of the embodiments of the present invention, rather than all of the embodiments. All other embodiments obtained by those of ordinary skill in the art based on the embodiments of the present invention without creative efforts shall fall within the protection scope of the present invention.

[0077] The purpose of the present invention is to provide a spacecraft attitude control and RCS optimal allocation method to solve the problems of low tracking control accuracy, large command error, and low control accuracy caused by large energy consumption of the existing spacecraft attitude within a limited time.

[0078] To make the above objects, features, and advantages of the present invention more obvious and understandable, the present invention will be further described in detail below in conjunction with the accompanying drawings and specific implementation manners.

[0079] Figure 1 and Figure 2 shows an exemplary process of the above spacecraft attitude control and RCS optimal allocation method, and the spacecraft attitude control and RCS optimal allocation method operates based on the predetermined performance of the spacecraft. Each step will be introduced in detail below.

[0080] Step 1: Obtain spacecraft parameters and spacecraft attitude control torque commands; the spacecraft parameters at least include the inertia matrix;

[0081] Step 2: Construct a spacecraft attitude dynamics model according to the spacecraft parameters, specifically including:

[0082]

[0083]

[0084] Among them, represents the roll in the inertial frame, ψ represents the pitch in the inertial frame, γ represents the yaw angle in the inertial frame, and θ ∈ R 3×1 represents the spacecraft attitude angle, represents the first derivative of θ, represents the first derivative of ω, and ω ∈ R 3×1 represents the spacecraft attitude angular velocity, and M is the three-axis torque acting on the spacecraft;

[0085] In one example, the inertia matrix is J.

[0086] Step 3: Construct a state-space equation according to the spacecraft attitude dynamics model; design a disturbance observer according to the state-space equation to observe and obtain an estimated value of the external disturbance torque acting on the spacecraft; specifically including:

[0087]

[0088]

[0089] A = -J -1 ω × J;

[0090] B = J -1 ;

[0091] Among them, x1 = θ, x2 = ω, A and B represent the equality sign, represents the first derivative of x2, represents the first derivative of x2, and u ∈ R 3×1 represents the control torque, and D ∈ R 3×1 represents the estimated value of the external disturbance torque acting on the spacecraft.

[0092] Design a disturbance observer according to the state-space equation to observe and obtain an estimated value of the external disturbance torque acting on the spacecraft, specifically including:

[0093]

[0094] Among them, L A = diag(L A1 L A2 L A3 ) > 0; k Ai = diag(k Ai1 k Ai2 k Ai3 ), i = 1,..., 4; κ Ai, α A are all predicted parameters; e x2 = x2 - x 2d ; x 1d represents the command value of the spacecraft attitude control torque command x1; x 2d represents the command value of the spacecraft attitude control torque command x2; the function sig r (x) = |x| r ·sgn(x);

[0095] The parameter σ = diag(σ1 σ2 σ3) is defined as follows:

[0096]

[0097] where, T Aj is a positive switching time parameter.

[0098] In one example, according to the state space equation, a uniformly convergent disturbance observer is designed, and the combined disturbance is estimated by the uniformly convergent disturbance observer

[0099] The parameter κ Ai is selected as:

[0100] κ A1 = 3I3κ A2 = 4.16I3κ A3 = 3.06I3κ A4 = 1.1I3;

[0101] where, the parameter k Ai needs to ensure that the polynomial s 4 + k A1j s 3 + k A2j s 2 + k A3j s + k A4j (j = 1, 2, 3) satisfies the Hurwitz condition, and the coefficient α A is a sufficiently small positive constant.

[0102]

[0103] where, T Aj is a positive switching time parameter; the estimation error can converge precisely within a fixed time which is the observed value of the designed disturbance observer.

[0104] Step 4: Construct a finite-time performance function according to the required control performance index; construct an error transformation state-space equation according to the finite-time performance function and the estimated value of the external disturbance torque received by the spacecraft; specifically including:

[0105] ρ fV (t) = a V3 t 4 + a V2 t 3 + a V1 t 2 + c Vρr t + c Vρ0 ;

[0106] Wherein, a V3 , a V2 and a V1 The expressions of are respectively:

[0107]

[0108] Wherein, the parameters c Vρ0 , c Vρr , ρ fV∞ and T fV are all predicted parameters, t represents time; c Vρ0 is a positive number, representing the initial error bound, c Vρr represents the initial change direction of the performance function, ρ fV∞ characterizes the steady-state convergence value of the function ρ fV (t), T fV represents ρ fV (t) converges to the steady-state value ρ fV∞ The set time of.

[0109] Construct an error transformation state-space equation according to the finite-time performance function and the estimated value of the external disturbance torque received by the spacecraft, specifically including:

[0110]

[0111]

[0112]

[0113] Wherein, ξ represents the transformation error, represents the first derivative of ξ, represents The first derivative of, Q and P represent the equality symbol, is the error upper limit coefficient of the finite-time performance function, and δ is the error lower limit coefficient of the finite-time performance function;

[0114] In one example, according to the finite-time performance function, the original error is converted into an unconstrained error, and the specific formula is:

[0115]

[0116] Step 5: Construct a super-twisting sliding mode controller according to the error conversion state space equation; specifically including:

[0117]

[0118]

[0119] where x'1 = ξ, represents the first derivative of x′1, represents the first derivative of x′2, represents the diagonal matrix composed of the vector Q, represents x 1d The second derivative of.

[0120] The sliding mode surface s in the super-twisting sliding mode controller is:

[0121] s = Cx'1 + x'2;

[0122] Super-twisting reaching law is:

[0123]

[0124]

[0125]

[0126] where, is the first derivative of s, C is the expected parameter, x′1, x′2 and v are equality symbols; K1, K2 are controller gains, is the first derivative of v; u d represents the control output command, represents the total disturbance observed by the disturbance observer;

[0127] According to the sliding mode surface, all sign functions in the reaching law are replaced with saturation functions to construct saturation functions, specifically including:

[0128]

[0129] where δ is a preset parameter.

[0130] Step 6: Construct a saturation function according to the sliding mode surface; construct a spacecraft attitude dynamics model with chattering suppression according to the super-twisting sliding mode controller and the saturation function;

[0131] In one example, a super-twisting sliding mode controller is designed for the above error conversion state space equation, and a saturation function is designed to replace the sign function to suppress chattering.

[0132] Step 7: Obtain the RCS optimal allocation result according to the spacecraft attitude control torque command and the spacecraft attitude dynamics model after suppressing chattering; the RCS optimal allocation result includes the on / off state of any one RCS thruster.

[0133] Obtaining the RCS optimal allocation result according to the spacecraft attitude control torque command and the spacecraft attitude dynamics model after suppressing chattering specifically includes:

[0134]

[0135] M RCS = F I R;

[0136] where M RCS represents the torque provided by the RCS, M d represents the spacecraft attitude control torque command, α is a weight coefficient, Q1 ∈ R 3×3 and Q2 ∈ R m×m are preset positive definite matrices; F I represents the coefficient matrix of the torque provided by the RCS, R = [r1, r2,..., r m T represents the on / off state of m RCS thrusters, and r i can only take 0 or 1.

[0137] In one example, the spacecraft control allocation problem is described as a quadratic programming problem according to the torque command, and the RCS optimal allocation is realized by the exhaustive method, and finally the on / off state of each RCS thruster is obtained.

[0138] For the above quadratic programming problem, the RCS allocation optimization problem is solved by the exhaustive method, and the specific steps are as follows:

[0139] First, calculate the values of the quadratic programming functions in all states of the RCS system, then compare the magnitudes of the function values in different states, and finally select the RCS system in the state with the smallest function value as the optimal optimal allocation result. According to the optimal allocation result, the states of each RCS thruster are obtained.

[0140] The RCS system of the spacecraft includes multiple thrusters. It is necessary to reasonably allocate the three-axis torque commands output by the designed super-twisting sliding mode controller to each thruster, and describe the RCS control allocation problem as a quadratic programming problem.

[0141] ​Among them, to further save the calculation time used by the exhaustive method, the quadratic programming function can be written as:

[0142]

[0143]

[0144]

[0145] L2 = αR T Q2R;

[0146] Among them, L0, L1, and L2 represent the pre-calculated quantities.

[0147] In other embodiments of the present invention, Matlab / Simulink is used for simulation to illustrate the specific calculation process of the embodiments of the present invention.

[0148] The simulation parameters are set as follows:

[0149] The spacecraft parameters are: inertia matrix The disturbance torque D = [500 + 100sin(πt / 125) 700 + 100cos(πt / 125) 400 + 100sin(πt / 125)] T , and the number of RCSs num = 16.

[0150] The predetermined performance parameters are: δ = [1 1 1] T , T fv = 10, c Vρ0 = 1, c Vρr = -0.02, ρ fV∞ = 0.01.

[0151] The parameters of the uniformly convergent observer are: T A = I 1×3 L A = 5I3, α A = 0.01, k A1j = 3, k A2j = 4.16, k A3j = 3.06, k A4j = 1.1.

[0152] The parameters of the super-twisting sliding mode controller are: C = 0.04I3, K1 = 1.5I3, K2 = 1.1I3, L0 = 0.1I3, δ = 0.005.

[0153] The RCS control allocation parameters are: Q1 = I3, Q2 = I 16 , σ = 0.1.

[0154] The simulation results without control allocation are as follows Figure 3 , Figure 4 , Figure 5 , Figure 6 , Figure 7 and Figure 8 shown below

[0155] From Figure 3 , Figure 4 , Figure 5 it can be seen that without considering control allocation (i.e., the actual torque value is equal to the commanded torque value), the designed super-twisting sliding mode controller can stably make the spacecraft attitude angle track the given commanded value, and the convergence time is less than the set time, and the tracking error is always within the set performance boundary

[0156] From Figure 6 , Figure 7 and Figure 8 it can be seen that due to no control allocation, the actual torque value is equal to the commanded torque value, so the two curves coincide, and the amplitude of curve chattering is very small, about 5 N·m

[0157] The simulation results under control allocation are as follows Figure 9 , Figure 10 , Figure 11 , Figure 12 , Figure 13 and Figure 14 shown below

[0158] From Figure 9 , Figure 10 , Figure 11 it can be seen that when considering control allocation, the designed super-twisting sliding mode controller can still stably make the spacecraft attitude angle track the given commanded value, and the convergence time is less than the set time, and the tracking error is always within the set performance boundary

[0159] From Figure 12 , Figure 13 and Figure 14 it can be seen that because control allocation is considered, the torque that the attitude control engine can provide is not a continuous quantity but a discrete quantity, so the actually provided torque value cannot be exactly equal to the commanded torque value, and the torque will also chatter due to this. Comparing Figure 6 , Figure 7 and Figure 8 it can be known that the chattering phenomenon is caused by control allocation, and the chattering of the controller itself is very small

[0160] In summary, in the embodiments of the present invention, a finite-time performance function is constructed, and an error transformation state-space equation is constructed based on the finite-time performance function and the estimated value of the external disturbance torque acting on the spacecraft. The control method with predetermined performance converts the original tracking error into an unconstrained error, thereby ultimately meeting some set control performance indicators, and can achieve precise tracking control of the spacecraft attitude within a finite time. According to the error transformation state-space equation, a super-twisting sliding mode controller is constructed; the super-twisting sliding mode control can effectively cope with large disturbances and will not cause serious chattering phenomena, and can simultaneously achieve the minimum command error and the comprehensive optimization of energy consumption control allocation optimization.

[0161] The various embodiments in this specification are described in a progressive manner. Each embodiment focuses on the differences from other embodiments. The same or similar parts among the various embodiments can be referred to each other. For the system disclosed in the embodiments, since it corresponds to the method disclosed in the embodiments, the description is relatively simple, and the relevant parts can be referred to the description of the method part.

[0162] Specific examples are used in this article to elaborate on the principles and implementation manners of the embodiments of the present invention. The description of the above embodiments is only used to help understand the method and its core idea of the embodiments of the present invention; at the same time, for those of ordinary skill in the art, based on the idea of the embodiments of the present invention, there will be changes in the specific implementation manners and application scopes. In summary, the content of this specification should not be construed as a limitation on the embodiments of the present invention.

Claims

1. A method for spacecraft attitude control and RCS optimal allocation, characterized in that, The spacecraft attitude control and RCS optimal allocation method operates based on the predetermined performance of the spacecraft, including: Obtaining spacecraft parameters and spacecraft attitude control torque commands; the spacecraft parameters at least include the inertia matrix; Constructing a spacecraft attitude dynamics model according to the spacecraft parameters; Constructing a state space equation according to the spacecraft attitude dynamics model; designing a disturbance observer according to the state space equation to obtain an estimated value of the external disturbance torque acting on the spacecraft; Constructing a finite-time performance function according to the required control performance index; constructing an error transformation state space equation according to the finite-time performance function and the estimated value of the external disturbance torque acting on the spacecraft; Constructing a super-twisting sliding mode controller according to the error transformation state space equation; Constructing a saturation function according to the sliding mode surface; constructing a spacecraft attitude dynamics model after chattering suppression according to the super-twisting sliding mode controller and the saturation function; Obtaining an RCS optimal allocation result according to the spacecraft attitude control torque command and the spacecraft attitude dynamics model after chattering suppression; the RCS optimal allocation result includes the on / off state of any RCS thruster.

2. The spacecraft attitude control and RCS optimal allocation method according to claim 1, wherein Constructing a spacecraft attitude dynamics model according to the spacecraft parameters, specifically including: ； ； Among them, , represents the roll in the inertial system, represents the pitch in the inertial system, represents the yaw angle in the inertial system, represents the spacecraft attitude angle, represents the first derivative of θ, represents the first derivative of, represents the spacecraft attitude angular velocity, is the three-axis torque acting on the spacecraft; Constructing a state space equation according to the spacecraft attitude dynamics model, specifically including: ； ； ； ； Among them, , , A and B represent the equality symbol, represents the first derivative of represents the first derivative of represents the control torque, represents the estimated value of the external disturbance torque acting on the spacecraft.

3. The spacecraft attitude control and RCS optimal allocation method according to claim 2, characterized in that, Designing a disturbance observer according to the state space equation to obtain an estimated value of the external disturbance torque acting on the spacecraft, specifically including: ； Among them, ; , ; , are all estimated parameters; ; represents the command value of the spacecraft attitude control torque command x1; represents the spacecraft attitude control torque command command value; function ; Parameter , is defined as follows: ； Among them, is the positive switching time parameter, is the time.

4. The spacecraft attitude control and RCS optimal allocation method according to claim 3, wherein Constructing a finite-time performance function according to the required control performance index, specifically including: ； Among them, and with The expressions are respectively: ； Among them, the parameters , , and are all predicted parameters, and t represents time; is a positive number, representing the initial error bound, represents the initial change direction of the performance function, characterizes the function steady-state convergence value, represents converges to the steady-state value settling time.

5. The spacecraft attitude control and RCS optimal allocation method according to claim 4, wherein Constructing an error transformation state space equation according to the finite-time performance function and the estimated value of the external disturbance torque acting on the spacecraft, specifically including: ； ； ； Among them, represents the conversion error, represents the first derivative of, , represents the first derivative of, Q and P represent the equality sign, is the upper limit coefficient of the finite-time performance function error, is the lower limit coefficient of the finite-time performance function error; Constructing a super-twisting sliding mode controller according to the error transformation state space equation, specifically including: ； ； Among them, , represents the first derivative of , represents the first derivative of represents the diagonal matrix formed by the vector , represents the second derivative of 6. The spacecraft attitude control and RCS optimal allocation method according to claim 5, wherein The sliding mode surface s in the super-twisting sliding mode controller is: ； Superhelical approaching law It is as follows: ； ； ； Among them, is the first derivative of s, is the predicted parameter, , and v are equality signs; K1 and K2 are controller gains, is the first derivative of v; represents the control output instruction, represents the total disturbance observed by the disturbance observer; Constructing a saturation function according to the sliding mode surface, specifically including: ； Among them, is a preset parameter.

7. The spacecraft attitude control and RCS optimal allocation method according to claim 6, wherein Obtaining an RCS optimal allocation result according to the spacecraft attitude control torque command and the spacecraft attitude dynamics model after chattering suppression, specifically including: ； ； Among them, represents the torque provided by the RCS, represents the spacecraft attitude control torque command, is the weight coefficient, and is a preset positive definite matrix; represents the coefficient matrix of the torque provided by the RCS, represents the on / off states of m RCS thrusters, can only take 0 or 1.

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