Attitude control energy optimization method for spacecraft

Abstract

The invention provides an attitude control energy optimization method for a spacecraft. The attitude control energy optimization method comprises the following steps: firstly, obtaining a smoothed attitude angle instruction and an attitude angle rate instruction by using a tracking differentiator; estimating impact disturbance through a nonlinear extended state observer; forming an impact disturbance sign signal according to the estimated value; obtaining a state error, establishing an attitude control law under a non-interference condition and an attitude control law under impact disturbance, and obtaining a corresponding control quantity and a sliding mode surface; according to the state error and the two attitude control laws, the switching compensation control quantity is calculated, the attitude control law under the impact disturbance containing integral switching compensation control is designed, and the smoothed control quantity under the impact disturbance is obtained; according to the impact disturbance mark signal, the control quantity under the interference-free condition or the control quantity under the smooth impact disturbance is selected and sent to the attitude control engine. The method effectively solves the energy optimization problem of the direct force control system when the parameters of the control system change drastically under large impact disturbance.

Description

TECHNICAL FIELD

[0001] The present application relates to the field of aircraft attitude control technology, in particular to an attitude control energy optimization method for a space vehicle. BACKGROUND

[0002] When a space vehicle flies outside the atmosphere, the control moment generated by the traditional aerodynamic rudder is significantly reduced due to the extremely thin air, and the control efficiency is poor. Therefore, a combined control mode of direct force and aerodynamic force is usually adopted. In this scheme, the orbit control engine is used to provide the required overload for the center of mass motion control of the vehicle, and the attitude control engine is responsible for generating the moment required to adjust the attitude motion. Since the space vehicle carries a limited amount of fuel for direct force control, the speed of fuel consumption directly depends on the design of the control method, so the energy use efficiency in the control process directly determines the on-orbit effective working time of the vehicle. The research on the energy optimization method for direct force control has important engineering application value for improving the mission performance.

[0003] At present, the energy optimization research for space vehicle direct force control mainly uses methods such as sequential quadratic programming and particle swarm algorithm to optimize the switch threshold of the thruster. This kind of method can suppress the frequent start-stop of the thruster caused by the dynamic change of the error by reasonably setting the switch threshold, so as to reduce fuel consumption and achieve the purpose of energy optimization. However, in actual application, especially in the face of large impact disturbance, this method has certain limitations. The existing optimization strategy fails to fully consider the dynamic switching behavior of the control system under strong disturbance and the additional energy consumption caused thereby, resulting in unsatisfactory energy optimization effect under large disturbance working conditions. SUMMARY

[0004] In view of the problems existing in the prior art, the present application proposes an attitude control energy optimization method for a space vehicle, which combines orbit control engine on-off impact disturbance estimation and smooth sliding mode disturbance-free switching control, effectively solving the problem of direct force control system energy optimization when the control system parameters change dramatically under large impact disturbance.

[0005] The application establishes a center of mass and attitude motion model of a space vehicle, a mass and center of mass change model, and an attitude control torque model; and establishes an impact disturbance model of on-off of a trajectory control engine by using a rectangular window function, estimates the impact disturbance of on-off of the trajectory control engine by using an extended state observer, and forms an impact disturbance signal; represents the attitude motion model of the space vehicle in a state space form, obtains a derivative of an attitude control command of the space vehicle by using a tracking differentiator, and designs a quasi-sliding mode attitude control law; designs different attitude control laws for no disturbance and impact disturbance respectively, further introduces a switching compensation control law for no disturbance switching control, establishes a bias model and a bias dynamic model of switching control; designs an integral sliding surface for the bias dynamic model of switching control, and designs a no disturbance switching compensation control law by using a smooth non-singular terminal sliding mode method.

[0006] Based on this, the technical scheme of the application is:

[0007] An attitude control energy optimization method of a space vehicle, comprising the following steps:

[0008] Step 1: obtaining an actual attitude angle and an attitude angle rate of the space vehicle, and a desired attitude angle command; processing the desired attitude angle command by using a tracking differentiator to obtain a smoothed attitude angle command and a smoothed attitude angle rate command;

[0009] Step 2: estimating an impact disturbance of on-off of a trajectory control engine of the space vehicle by using a nonlinear extended state observer according to system states and control inputs, to obtain an estimated value of the impact disturbance;

[0010] Step 3: forming a binary impact disturbance flag signal by using a disturbance threshold according to the estimated value of the impact disturbance;

[0011] Step 4: obtaining a state error according to system states and the smoothed attitude angle command and the smoothed attitude angle rate command, and establishing an attitude control law under no disturbance condition and an attitude control law under impact disturbance condition, to obtain a control quantity under no disturbance condition and a sliding surface, and a control quantity under impact disturbance condition and a sliding surface;

[0012] Step 5: calculating a switching compensation control quantity according to the state error in step 4, the attitude control law under no disturbance condition and the attitude control law under impact disturbance condition; designing an attitude control law under impact disturbance condition containing integral switching compensation control by using the switching compensation control quantity, to obtain a smoothed control quantity under impact disturbance condition;

[0013] Step 6: selecting the control quantity under no disturbance condition or the smoothed control quantity under impact disturbance condition according to the impact disturbance flag signal and sending it to an attitude control engine, to realize energy optimal attitude control.

[0014] Further preferred, in step 1, the tracking differentiator is

[0015]

[0016] wherein, is the input attitude angle command, is the smoothed attitude angle command, is the derivative of , is the velocity factor, is the calculation step size; is the optimal control quantity, is the bang-bang control function.

[0017] Further preferred, in step 1, the bang-bang control function is

[0018]

[0019] wherein is the step size dependent quantity; is the final synthesized control quantity, which is switched between and the candidate optimal control quantity according to the magnitude of the mixing error , to form the final control quantity ; is the sign function; the specific expression is:

[0020] .

[0021] Further preferred, in step 2, the nonlinear extended state observer is

[0022]

[0023] wherein, , are the estimated values of the state quantities and , respectively, is the estimated value of the external disturbance , , , are the observer gain parameters, the state quantity is the attitude angle, the state quantity is the attitude angle rate, is the known part in the space vehicle dynamics model, is the control gain, is the control input, the function is calculated as follows:

[0024]

[0025] wherein is a sign function, is a nonlinear factor, is a filter factor.

[0026] In a further preferred embodiment, in step 3, the impact disturbance flag signal is

[0027]

[0028] wherein is a disturbance threshold, the impact disturbance flag signal takes the value 0 for no disturbance and the value 1 for a disturbance.

[0029] In a further preferred embodiment, in step 4, the attitude control law under no disturbance conditions is

[0030]

[0031] wherein is the control quantity under no disturbance conditions, is the sliding surface under no disturbance conditions, is the rated torque provided by the attitude control engine, , is the sliding surface coefficient under no disturbance conditions, is a saturation function, is a quasi-sliding mode control threshold;

[0032] the attitude control law under impact disturbance conditions is

[0033]

[0034] wherein is the control quantity under impact disturbance conditions, is the sliding surface under impact disturbance conditions, , is the sliding surface coefficient under impact disturbance conditions.

[0035] In a further preferred embodiment, in step 5, the switching compensation control quantity is

[0036]

[0037] wherein is the state matrix of the deviation system, is the input matrix of the deviation system, is the disturbance input matrix of the deviation system, is the sliding surface switching deviation, is the integral sliding mode surface coefficient, is the set terminal attractor index, , , and is the convergence characteristic adjustment parameter, is the integral sliding mode surface, .

[0038] In a further preferred embodiment, in step 5, the sliding mode surface switching deviation is:

[0039] e ( t ) = s ˜ b − s a = E [ e x , e ˙ x ] T + k c ∫ 0 t u c

[0040] wherein is the sliding mode surface under the smoothed impact disturbance, , is the attitude angle control error, is the switching compensation gain.

[0041] In a further preferred embodiment, in step 5, the integral sliding mode surface is:

[0042] .

[0043] In a further preferred embodiment, in step 5, the attitude control law under the impact disturbance containing the integral type switching compensation control is:

[0044] .

[0045] Advantages

[0046] The advantages of the present application mainly manifest in the following two aspects:

[0047] Firstly, for the two typical working conditions of no disturbance and strong impact disturbance, attitude control laws with different parameter configurations are respectively designed. The hierarchical control strategy can maintain the required attitude control accuracy of the system under different disturbance intensities, effectively enhancing the adaptability and robustness of the control system to the external disturbance environment.

[0048] Secondly, the smooth non-singular terminal sliding mode method is introduced to construct a smooth switching control law suitable for different disturbance states. This method can realize smooth transition of the control system between different modes, avoid control chattering or oscillation problems caused by hard switching, thereby significantly reducing the frequent start-stop phenomenon of the attitude control engine in the switching process, which is conducive to energy optimization of the attitude control system under impact disturbance and improves fuel efficiency.

[0049] Additional aspects and advantages of the application will be set forth in part in the following description, will become apparent from the following description, or will be learned through practice of the application. BRIEF DESCRIPTION OF DRAWINGS

[0050] The above and / or additional aspects and advantages of the present application will become apparent and be readily appreciated from the following description, including the references to the figures, in which:

[0051] Figure 1 General design idea for embodiments of the present application

[0052] Figure 2 Trajectory control impact disturbance estimation graph for Example 1

[0053] Figure 3 Attitude control simulation graph for hard switching in Example 1

[0054] Figure 4 Attitude control simulation graph for disturbance-free switching in Example 1

[0055] Figure 5 Attitude control fuel consumption graph for hard switching in Example 1

[0056] Figure 6 Attitude control fuel consumption graph for disturbance-free switching in Example 1

[0057] Figure 7 Attitude control simulation graph for hard switching in Example 2

[0058] Figure 8 Attitude control simulation graph for disturbance-free switching in Example 2

[0059] Figure 9 Attitude control fuel consumption graph for hard switching in Example 2

[0060] Figure 10 Attitude control fuel consumption graph for disturbance-free switching in Example 2

[0061] Figure 11 Attitude control simulation graph for hard switching in Example 3

[0062] Figure 12 Attitude control simulation graph for disturbance-free switching in Example 3

[0063] Figure 13 Attitude control fuel consumption graph for hard switching in Example 3

[0064] Figure 14 Attitude control fuel consumption graph for disturbance-free switching in Example 3 DETAILED DESCRIPTION

[0065] Embodiments of the present application are described in detail below, which are exemplary and intended to explain the present application, and should not be understood as a limitation of the present application.

[0066] This invention provides an energy optimization method for attitude control of a spacecraft. In this embodiment, the spacecraft is an exoatmospheric interceptor. This method solves the problems of insufficient optimization capability and difficulty in guaranteeing accuracy of existing direct force attitude control system energy optimization methods under transient impact disturbances. It can effectively improve the adaptability of the interceptor's attitude control system to transient impact disturbances.

[0067] The technical solution in this application example addresses the attitude control energy optimization problem under the aforementioned instantaneous impact disturbance conditions. The overall approach is as follows: Figure 1 As shown, it includes the following:

[0068] (i) Establish the center of mass motion model, attitude kinematics and dynamics model, and attitude control torque model of the interceptor;

[0069] The equation of motion for the interceptor's center of mass can be expressed as:

[0070] (1)

[0071] In the formula, , , These are the position components of the interceptor in the launch coordinate system. For the interceptor speed, The trajectory angle of the interceptor. This refers to the trajectory deflection angle of the interceptor.

[0072] The attitude kinematic equations of the interceptor can be expressed as:

[0073] (2)

[0074] In the formula, , , are the attitude angles of the interceptor in the launch coordinate system, where are... The pitch angle, Yaw angle For roll angle, , , These are the interceptor's attitude angular velocities in the projectile coordinate system; where... For the roll rate, The yaw rate, denoted as pitch rate.

[0075] The attitude dynamics equations of the interceptor can be expressed as:

[0076] (3)

[0077] In the formula, , , are the rotational inertia of the interceptor relative to the body coordinate system, , , are the roll, yaw and pitch control moments of the interceptor.

[0078] According to the layout of the attitude control engine in the interceptor, the control moments of the three channels are expressed as:

[0079] (4)

[0080] where, , , are the distances between the thrusts of the attitude control engines of the three channels and the center of mass of the interceptor; , is the diameter of the interceptor.

[0081] (II) Establish an impact disturbance model of the interceptor and design an impact disturbance signal;

[0082] (1) Modeling of the impact disturbance of the on-off of the orbit control engine of the interceptor;

[0083] By comparing the on-off command of the orbit control engine of the interceptor and the test data of the disturbance moment, the response time, the disturbance moment amplitude, the rising edge and the falling edge time of the orbit control direct force system are analyzed, and the impact disturbance model of the on-off of the orbit control engine is established based on the rectangular window function:

[0084] (5)

[0085] where, is the disturbance moment at time t, which represents the instantaneous interference moment applied to the interceptor due to the on-off of the orbit control engine; is a time-varying quantity, and its variation law is simplified as a rectangular pulse, is the impact disturbance amplitude, which is determined by analyzing the actual test data of the orbit control engine, is the rectangular window function used to simulate the instantaneous pulse form of the on-off of the orbit control engine, is the duration of the impact disturbance, is the response delay time, which represents the time difference between the time when the on-off command of the orbit control engine is issued by the control system and the time when the disturbance moment actually acts on the spacecraft. This delay is caused by the physical processes such as the action of the fuel valve and the establishment of the pipeline pressure.

[0086] ​(2) The same method is used to design a nonlinear extended state observer for each channel to estimate the impact disturbance of the on-off of the orbit control engine.

[0087] Taking the pitch channel as an example, represents the pitch angle , represents the pitch angle rate , and the interceptor attitude motion state equation of the pitch channel can be expressed as:

[0088] (6)

[0089] In the formula, is the known part in the dynamic model of the system, that is, the known part of the interceptor attitude dynamics equation, except for the disturbance, is the control gain, is the control input;

[0090] The nonlinear extended state observer is designed as follows:

[0091] (7)

[0092] In the formula, , are the estimated values of the state variables and , respectively, is the estimated value of the external disturbance , , , are the observer gain parameters. The function is calculated as follows:

[0093] (8)

[0094] In the formula, is the sign function, the nonlinear factor and the filtering factor are used to balance the convergence speed and smoothness of the observer.

[0095] (3) According to the estimated value of the external disturbance , an impact disturbance flag signal is formed by designing a disturbance threshold:

[0096] (9)

[0097] In the formula, is the disturbance threshold, and the impact disturbance flag signal The value of 0 represents no interference, 1 represents the existence of interference, and is used to intelligently decide which set of control strategies to enable according to the current working condition.

[0098] (Three) use the formed impact disturbance flag signal to design a non-disturbance switching control algorithm.

[0099] (1) For each channel, the same method is used to design a quasi-sliding mode attitude control system based on a tracking differentiator;

[0100] Take the pitch channel as an example:

[0101] First, the pitch channel attitude motion equation of the interceptor is expressed in the form of the following state space equation:

[0102] (10)

[0103] In the formula, is the system state vector, is the system matrix, is the input matrix, is the output matrix, is the disturbance input matrix, is the output vector of the system, is the control input of the system, is the switching signal of the controller, is the external disturbance signal.

[0104] Secondly, the quasi-sliding mode control model of the pitch channel is designed:

[0105] The sliding mode of the pitch channel is designed:

[0106] (11)

[0107] In the formula, is the pitch channel sliding surface, is the pitch angle command, , are the pitch angle and pitch angle rate of the interceptor, , is the pitch channel sliding surface coefficient.

[0108] The first-order derivative of the pitch angle command in the above formula is calculated by the following tracking differentiator:

[0109] (12)

[0110] In the formula, is the input pitch angle command ; and are the pitch angle command The tracking signal and its first derivative The tracking signal As a smoothed pitch angle command, if the input pitch angle command... A step transition occurs. It will transition to the new instruction value with a smooth, overshoot-free curve. yes The differential; For velocity factor, To calculate the step size. This is the optimal control quantity. The steepest control function is calculated as follows:

[0111] (13)

[0112] In the formula, This is a step-size related quantity; The final synthesized control quantity is based on the mixing error. The size is determined by the switching function. exist and candidate optimal control quantity Smooth switching between them to form the final control quantity. ; It is a symbolic function; its specific expression is:

[0113] (14)

[0114] The quasi-sliding mode control model for the pitch channel is obtained as follows:

[0115] (15)

[0116] In the formula, The rated torque that the attitude control engine can provide. The threshold for sliding mode control.

[0117] (2) Using the same method, design the control system parameters for each channel under both interference-free and interference-affected conditions;

[0118] Taking the pitch channel as an example:

[0119] First, design the interceptor pitch attitude control law under interference-free conditions, i.e. hour:

[0120] (16)

[0121] The control quantity under no-interference conditions, For the sliding surface under undisturbed conditions, , For the sliding surface under impact disturbance, , Usually set relatively mild to maintain smooth control and reduce fuel consumption; For the saturation function;

[0122] Next, the interceptor pitch attitude control law under impact disturbance is designed, that is, When:

[0123] (17)

[0124] For the control under impact disturbance, For the sliding surface under impact disturbance, , For the sliding surface coefficient under impact disturbance, , Usually set more aggressive (e.g. larger gain) to enable the system to quickly overcome disturbances.

[0125] (3) When Changes, directly switch And Will cause the control force to jump, so the smooth nonsingular terminal sliding mode based disturbance-free switching control law is designed for each channel respectively;

[0126] Take the pitch channel as an example:

[0127] On the basis of the pitch attitude control law under impact disturbance in (2), further design the pitch attitude control law under impact disturbance containing integral type switching compensation control:

[0128] (18)

[0129] Where For the smooth control under impact disturbance, For the smooth sliding surface under impact disturbance, For the switching compensation control, For the switching compensation gain, used to adjust the influence of switching compensation control On the sliding surface;

[0130] Switching compensation control Obtained by the following process:

[0131] According to the pitch attitude control law containing integral type switching compensation control and the pitch attitude control law under disturbance-free condition, the pitch attitude control law deviation model is constructed:

[0132] e ( t ) = s ˜ b − s a = E [ e x , e ˙ x ] T + k c ∫ 0 t u c (19)

[0133] wherein, is the sliding mode surface switching error, is the pitch angle control error, is the two kinds of control law parameter error matrix.

[0134] Next, the pitch attitude control law error model is taken one order differential to establish the pitch attitude error dynamic model:

[0135] (20)

[0136] wherein, is the state matrix of the error system, is the generalized inverse matrix of matrix , is the input matrix of the error system, is the disturbance input matrix of the error system;

[0137] Based on the established pitch attitude error dynamic model, the integral sliding mode surface is designed as follows:

[0138] (21)

[0139] wherein, is the integral sliding mode surface coefficient, , is the set terminal attractor index;

[0140] Based on the smooth nonsingular terminal sliding mode, the disturbance-free switching compensation control law of the pitch channel is designed as follows:

[0141] (22)

[0142] wherein, is the feedforward compensation of the state error, is the feedforward compensation of the control input, is the feedforward compensation of the disturbance, , and are the convergence characteristic adjustment parameters.

[0143] According to the above, the attitude control energy optimization method of the interceptor specifically includes the following steps:

[0144] Step 1: Obtain the actual attitude angle and attitude angle rate of the interceptor, and the expected attitude angle command; for example, the actual attitude angle and attitude angle rate of the interceptor are obtained through gyroscopes, star sensors and other sensors, and the expected attitude angle command is obtained through a guidance system; the attitude angle includes the pitch angle , the yaw angle and the roll angle Attitude angular rate includes roll angular rate yaw rate pitch rate Since the methods for pitch, roll, and yaw are the same, we will take pitch angle and pitch rate as an example here, that is, to obtain the actual pitch angle of the interceptor. and pitch rate and desired pitch angle command .

[0145] The desired attitude angle command is processed using a tracking differentiator, i.e., the smoothed attitude angle command is obtained according to formulas (12)-(14). and the smoothed attitude angular rate command .

[0146] Step 2: Based on the system state and control input, estimate the impact disturbance of the track-controlled engine start-up and shutdown using a nonlinear extended state observer to obtain the estimated value of the impact disturbance. Here, the system state refers to the actual attitude angle and attitude angular rate of the interceptor, while the control input is the controller output obtained at the end of the previous control cycle. The initial value of the control input can be set to 0 or a very small value. Similarly, since the methods for pitch, roll, and yaw channels are the same, we will take the pitch channel as an example, based on the actual pitch angle of the interceptor... and pitch rate and control input The estimated value of the shock disturbance is obtained by using the nonlinear extended state observer, i.e., formulas (7)-(8). .

[0147] Step 3: Based on the estimated value of the impact disturbance, a binary impact disturbance flag signal is generated using the disturbance threshold; as shown in formula (9), based on the estimated value... The impact disturbance signal was obtained. .

[0148] Step 4: Based on the system state, the smoothed attitude angle command, and the smoothed attitude angular rate command, obtain the state error, and establish the attitude control law under no-disturbance conditions and the attitude control law under impact disturbance, preparing the optimal control strategy for different operating conditions. Taking the pitch channel as an example, according to formulas (16)-(17), the pitch attitude control law under no-disturbance conditions and the pitch attitude control law under impact disturbance are designed, and the control quantity under no-disturbance conditions is obtained. and sliding surfaces Control quantity under impact disturbance and sliding surfaces ;

[0149] Step 5: According to the state error in step 4, the attitude control law under the condition of no disturbance and the attitude control law under the condition of impact disturbance, the switching compensation control amount is calculated; the switching compensation control amount is used to design the attitude control law under the condition of impact disturbance containing integral type switching compensation control. Still taking the pitch channel as an example, according to formulas (19)-(22), the switching compensation control amount is obtained , which is substituted into formula (18) to design the attitude control law under the condition of impact disturbance containing integral type switching compensation control, and the smooth control amount under the condition of impact disturbance is obtained . .

[0150] Step 6: According to the impact disturbance flag signal , the control amount under the condition of no disturbance or the smooth control amount under the condition of impact disturbance is sent to the attitude control engine to realize energy-optimal attitude control.

[0151] The following gives a specific embodiment:

[0152] Embodiment 1:

[0153] The attitude control of the interceptor under the condition of linear attitude angle command is taken as an example to implement the method. The initial pitch angle and angular velocity of the interceptor are zero, the total simulation time is 10 seconds, and the impact disturbance occurs in the 2-4 seconds. Figure 2 is the estimation result of the disturbance observer for judging the impact disturbance of the orbit control.

[0154] Figure 3 and Figure 4 are the attitude angle tracking control simulation results of the interceptor under the hard switching and disturbance-free switching of the control system, respectively. It can be seen that the attitude response under the disturbance-free switching is more smooth;

[0155] Figure 5 and Figure 6 are the fuel consumption simulation results of the interceptor under the hard switching and disturbance-free switching of the control system, respectively, which are 0.231 Kg and 0.186 Kg, respectively. It can be seen that the fuel consumption under the disturbance-free switching is less;

[0156] Embodiment 2:

[0157] The attitude control of the interceptor under the condition of sinusoidal attitude angle command is taken as an example to implement the method. The initial pitch angle and angular velocity of the interceptor are zero, the total simulation time is 10 seconds, and the impact disturbance occurs in the 2-4 seconds.

[0158] Figure 7 and Figure 8 are the attitude angle tracking control simulation results of the interceptor under the hard switching and disturbance-free switching of the control system, respectively. It can be seen that the sinusoidal command tracking curve under the disturbance-free switching is more stable;

[0159] Figure 9 and Figure 10 are respectively the fuel consumption simulation results of the interceptor under the hard switching and the bumpless switching of the control system, and are respectively 0.238 Kg and 0.223 Kg, it can be seen that the fuel consumption under the bumpless switching is less;

[0160] Embodiment 3:

[0161] The method is implemented by taking the attitude control of the interceptor under the square wave type attitude angle command as an example, the initial pitch angle and the angular velocity of the interceptor are zero, the total simulation time is 10 seconds, and the impact disturbance occurs in the 2-4 seconds.

[0162] Figure 11 and Figure 12 are respectively the attitude angle tracking control simulation results of the interceptor under the hard switching and the bumpless switching of the control system, it can be seen that the steady-state error of the square wave command tracking under the bumpless switching is smaller;

[0163] Figure 13 and Figure 14 are respectively the fuel consumption simulation results of the interceptor under the hard switching and the bumpless switching of the control system, and are respectively 0.575 Kg and 0.529 Kg, it can be seen that the fuel consumption under the bumpless switching is less.

[0164] Although the embodiments of the present application have been shown and described above, it can be understood that the above-mentioned embodiments are exemplary and cannot be understood as limiting the present application, and those skilled in the art can make changes, modifications, replacements and variations to the above-mentioned embodiments without departing from the principles and purposes of the present application within the scope of the present application.

Claims

1. A method for attitude control energy optimization of a space vehicle, characterized by: The method comprises the following steps: Step 1: obtaining the actual attitude angle and attitude angle rate of a space vehicle, and a desired attitude angle instruction; processing the desired attitude angle instruction by using a tracking differentiator to obtain a smoothed attitude angle instruction and a smoothed attitude angle rate instruction; Step 2: estimating the impact disturbance of the on-off of the orbit control engine of the space vehicle according to the system state and the control input by using a nonlinear extended state observer to obtain an estimated value of the impact disturbance; Step 3: forming a binary impact disturbance flag signal according to the estimated value of the impact disturbance by using a disturbance threshold; Step 4: obtaining a state error according to the system state and the smoothed attitude angle instruction and the smoothed attitude angle rate instruction, and establishing an attitude control law under a non-disturbance condition and an attitude control law under an impact disturbance condition to obtain a control quantity under the non-disturbance condition and a sliding surface, a control quantity under the impact disturbance condition and a sliding surface; Step 5: calculating a switching compensation control quantity according to the state error in Step 4 and the attitude control law under the non-disturbance condition and the attitude control law under the impact disturbance condition; designing an attitude control law under the impact disturbance condition containing integral switching compensation control by using the switching compensation control quantity to obtain a smoothed control quantity under the impact disturbance condition; Step 6: selecting the control quantity under the non-disturbance condition or the smoothed control quantity under the impact disturbance condition according to the impact disturbance flag signal and sending the selected control quantity to the attitude control engine to realize energy-optimal attitude control.

2. The method of claim 1, wherein: In Step 1, the tracking differentiator is: wherein is the input attitude angle command, is the smoothed attitude angle command, is the derivative of is the velocity factor, is the computation step size; is the optimal control amount, is the bang-bang control function.​ 3. The method of claim 2, wherein: In Step 1, the fastest control function is wherein is a step related quantity; is a final synthesis control quantity, which is switched between and a candidate optimal control quantity according to the magnitude of the mixing error and forms the final control quantity ; ; is a sign function; the specific expression is: 。 4. The method of claim 2, wherein: In Step 2, the nonlinear extended state observer is: where , are the estimates of the state variables and , respectively, is the estimate of the external disturbance , and , , are the observer gain parameters, the state variables is the attitude angle, the state variable is the attitude angular rate, is the known part of the spacecraft dynamics model, is the control gain, is the control input, the function is calculated as follows: wherein is a symbol function, is a non-linear factor, is a filter factor.

5. The method of claim 4, wherein: In Step 3, the impact disturbance flag signal is: wherein is a disturbance threshold, the impulse disturbance flag signal a value of 0 indicates no disturbance, a value of 1 indicates a disturbance.

6. The method of claim 4, wherein: In Step 4, the attitude control law under the non-disturbance condition is: wherein is the control variable under the condition of no disturbance, is the sliding mode surface under the condition of no disturbance, is the rated torque size that the attitude control engine can provide, , is the sliding mode surface coefficient under the condition of no disturbance, is a saturation function, is a quasi-sliding mode control threshold; The attitude control law under the impact disturbance condition is: wherein is the control quantity under the impact disturbance, is the sliding mode surface under the impact disturbance, , is the sliding mode surface coefficient under the impact disturbance.

7. The method of claim 6, wherein: In Step 5, the switching compensation control quantity is: wherein is a state matrix of the bias system, is an input matrix of the bias system, is a disturbance input matrix of the bias system, is a sliding surface switching bias, is an integral sliding surface coefficient, is a set terminal attractor index, , , and is a convergence characteristic adjustment parameter, is an integral sliding surface, .

8. The method of claim 7, wherein: In Step 5, the switching deviation of the sliding surface is: wherein is the sliding mode surface under the smoothed impact disturbance, , is the attitude angle control error, is the switching compensation gain.

9. The method of claim 7, wherein: In Step 5, the integral sliding surface is: 。 10. The method of claim 7, wherein: In Step 5, the attitude control law under the impact disturbance condition containing integral switching compensation control is: 。

Images