CN115837989B - On-orbit target approach guidance method based on attitude-orbit coupling control strategy - Google Patents
On-orbit target approach guidance method based on attitude-orbit coupling control strategy Download PDFInfo
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Abstract
The invention discloses an on-orbit target approach guidance method based on an attitude and orbit coupling control strategy, which comprises the steps of establishing a rigid body attitude angle dynamic model of a spacecraft to obtain a functional relation of yaw angular acceleration and pitch angular acceleration of the spacecraft with respect to cold air spraying force of a horizontal plane and a vertical plane under an inertial coordinate system and a nonlinear functional change relation of yaw angular velocity and pitch angular velocity and time of the spacecraft; the attitude orbit correction control strategy is established, so that the relative sight angle is ensured to be in the center of the view field of the spacecraft and the accuracy of the off-target quantity is ensured; introducing auxiliary variables related to attitude angles of the spacecraft to correct the command acceleration of a plane where a yaw angle of the spacecraft is positioned and the command acceleration of a plane where a pitch angle of the spacecraft is positioned in a generalized proportional guidance law; and combining with a Schmitt trigger logic, constructing a PWM pulse ignition strategy, and enabling the spacecraft to carry out track correction according to an improved pulse generalized proportion guiding method so as to achieve high-precision off-target quantity. The invention enables the target to be positioned in the central area of the field of view, and simultaneously reduces the problem of gesture rail dynamic coupling.
Description
Technical Field
The invention designs a space flight and aviation technology, and particularly relates to an on-orbit target approach guidance method based on an attitude and orbit coupling control strategy.
Background
With the deep penetration of human exploration, development and utilization of outer space and the demand of on-orbit services of spacecraft, higher requirements are put on the meeting docking technology. Problems such as on-orbit capture and maintenance of a faulty spacecraft, space debris removal, changing of the asteroid orbit and space defence have become the subject to be faced and solved by the development of aerospace technology, while approximation technology to non-cooperative targets is a critical technology necessary to solve these problems. Spatially non-cooperative targets generally refer to a class of spatial objects that cannot provide valid information, including failed or failed satellites, spatial debris, asteroid, and opposing spacecraft, among others. The approximation process involves relative movements of the spacecraft, including local movements of the controlled individual and coupled movements of the spacecraft attitude trajectories. In order to meet the approaching requirement of the spacecraft on the non-cooperative target, the on-orbit target approaching tail end guidance law technology based on the attitude and orbit coupling problem is researched and has important value.
The on-orbit target approximation process can be roughly divided into a long-distance guiding section, a middle-distance approaching section and a short-distance end approaching section according to the relative distance between the controlled spacecraft and the non-cooperative target, wherein the short-distance end approaching section starts from a position with a distance of hundreds of meters to thousands of meters from the target, and is a core stage of the on-orbit approximation stage, and is widely focused by related researchers. The on-orbit target short-distance approximation segment not only needs to design the relative position between the controlled spacecraft and the target controlled by guidance law, but also needs to study the attitude control problem of the spacecraft, including the study of an attitude orbit control strategy.
Aiming at a high-speed moving target, the proportional guidance law is widely applied to various guidance tasks, in engineering practice, the proportional guidance law can be divided into an amplified proportional guidance law and a generalized proportional guidance law after improvement, and a proper proportional guidance law can be selected according to actual requirements, but the proportional guidance law only aims at track movement, both a spacecraft and the target are regarded as particles, and the influence of gesture change on track guidance is not considered. Peng Qian et al propose a high-precision and low-energy-consumption attitude and orbit integrated composite control method for an adjacent space interceptor, design a mapping function containing line-of-sight angle constraint by utilizing an arctangent function, and design an attitude and orbit integrated composite control law by combining high-order sliding mode control aiming at the attitude and orbit coupling problem. Liu Xiaoma et al design guidance laws by using variable structure control theory and optimal control theory for the two cases of larger target maneuver and smaller target maneuver, and design attitude and orbit control laws based on the pulse working mode of the attitude and orbit control engine. The guidance law and the attitude control law are based on the premise of dual-channel control of the spacecraft, the orbit control engine is positioned at the centroid position, the attitude control engine is symmetrical to the centroid, and the centroid position of the spacecraft is occupied in the study, so that the attitude correction is realized only by the engine symmetrical to the centroid based on attitude and orbit coupling. Wu Fan and the like invent a gesture-rail integrated control method based on pulse width modulation of rail-controlled thrusters, and selects a thruster switching signal by using gesture angle deviation based on the layout that 4 thrusters are positioned on an XOY plane. However, the design problem of the guidance law is not considered in the invention, and only the control strategy of the two-dimensional plane is considered. Zheng Huixin et al design a variable coefficient guidance law based on an augmentation ratio guidance law, and design an ignition strategy to convert a continuous instruction into a pulse engine instruction in consideration of the motion acceleration of a target, aiming at the characteristics of high relative speed, high impact precision requirement and limited fuel of a high-speed impact task. However, in this study, the influence of the attitude change on the guidance law command was not considered, and the command control period and the restriction of the thrust magnitude were not considered.
In fact, in engineering practical application, not only the problem of attitude and orbit coupling, but also the situation that an attitude and orbit engine is not at the mass center position are considered, and meanwhile, the constraint of a control period and the constraint of the thrust magnitude direction exist when a control instruction is output. Therefore, based on the situation, the invention combines the time-sharing multiplexing theory to design the attitude and orbit control strategy and combines the improved pulse generalized proportion guidance law to control the spacecraft, so that the attitude and orbit correction is carried out in the process of approaching the target, and finally the high-precision off-target quantity is achieved.
Disclosure of Invention
The invention mainly aims to provide an on-orbit target approach guidance method based on a gesture-orbit coupling control strategy.
The technical scheme for realizing the aim of the invention is as follows: an on-orbit target approach guidance method based on an attitude-orbit coupling control strategy comprises the following steps:
step 1: establishing a spacecraft attitude dynamics model
According to the structural layout of the spacecraft air nozzle symmetrical about the centroid and the rotation dynamics equation of the longitudinal axis around the centroid, a spacecraft rigid body attitude angle dynamic model is established, and the yaw angular acceleration of the spacecraft is obtainedIs +.>Regarding the cold air spraying force F of the horizontal plane and the vertical plane under the inertial coordinate system a 、F b Is a function of the yaw rate w of the spacecraft a Angular velocity w from pitch angle b A nonlinear function change relation with time t;
step 2: construction of attitude and orbit correction control strategy aiming at attitude and orbit correction coupling problem
According to yaw angular acceleration of spacecraftIs +.>Leng Qipen force F with respect to horizontal and vertical planes in inertial frame a 、F b Is a function of the yaw rate w of the spacecraft a Angular velocity w from pitch angle b Constructing a pose rail correction control strategy according to the nonlinear function change relation of time t, and ensuring that the relative sight angle is in the center of the view field of the spacecraft;
step 3: introducing auxiliary variable related to attitude angle of spacecraft to improve generalized proportional guidance law
Introducing an auxiliary variable epsilon with respect to attitude angle of spacecraft a 、ε b Correcting instruction acceleration a of plane where yaw angle of spacecraft in generalized proportional guidance law is located a Commanded acceleration a of plane in which pitch angle is located b To reduce the error of the instruction acceleration caused by the existence of the attitude angle;
step 4: PWM pulse ignition strategy design according to improved generalized proportional pilot law
According to the plane instruction acceleration a of the yaw angle of the spacecraft a The plane command acceleration a of pitch angle b Regarding the continuous time-varying function change relation of time t, a Pulse Width Modulation (PWM) pulse ignition strategy is constructed by combining with a Schmitt trigger logic, so that a spacecraft enters according to an improved pulse generalized proportion guiding methodAnd correcting the line track to achieve high-precision off-target quantity.
An on-orbit target approach guidance system based on an attitude-orbit coupling control strategy realizes on-orbit target approach guidance based on the attitude-orbit coupling control strategy based on the on-orbit target approach guidance method.
A computer device comprises a memory, a processor and a computer program stored on the memory and capable of running on the processor, wherein when the processor executes the computer program, on the basis of the on-orbit target approach guidance method, on-orbit target approach guidance based on a gesture-orbit coupling control strategy is realized.
A computer readable storage medium having stored thereon a computer program which, when executed by a processor, implements on-orbit target approach guidance based on an on-orbit target approach guidance method described.
Compared with the prior art, the method considers the design of the tail end guidance law of the three-dimensional space in the process of approaching the on-orbit target of the spacecraft, designs a gesture track control strategy for weakening the gesture track coupling problem, and considers that not only gesture adjustment but also track correction is completed in the whole tail end guidance process, thereby achieving high-precision off-target quantity and improving the time utilization rate. And the attitude angle factor and the control period factor are introduced in the guidance law design, so that the guidance law of the existing missing distance measurement information is more in line with the actual operation law of the spacecraft and is more close to engineering application. The method can be widely applied to satellite recovery tasks of military operations of spacecraft approaching on-orbit non-cooperative targets.
Drawings
FIG. 1 is a flow chart of an on-orbit target near-end guidance law design method algorithm based on a gesture-orbit coupling control strategy.
Fig. 2 is a flow chart of the attitude and orbit correction control strategy.
Fig. 3 is a relationship of the change of the attitude angular velocity of the primary attitude orbit correction spacecraft with time t.
FIG. 4 is a schematic diagram of relative line of sight of a spacecraft and a target, and a spacecraft generalized proportional lead command acceleration versus a spacecraft longitudinal axis.
Fig. 5 is a graph of the geometry of a spacecraft with a target.
Fig. 6 is a three-dimensional trajectory of a spacecraft with a target.
FIG. 7 is a graph of relative distance between a spacecraft and a target over time.
Fig. 8 shows the relationship of the time-dependent cold air correction force in the horizontal direction in the inertial coordinate system.
Fig. 9 shows the relationship of the time-dependent cold air correction force in the vertical direction in the inertial coordinate system.
Fig. 10 is a time-dependent relationship of the deviation between the yaw angle and the azimuth angle of the line of sight of the spacecraft.
FIG. 11 is a graph of the deviation between pitch angle and line of sight angle of a spacecraft as a function of time.
Fig. 12 is a time-dependent relationship of the yaw angle and rotational angular velocity of the spacecraft.
Fig. 13 is a graph showing the relationship of the pitch angle rotation angular velocity of the spacecraft with time.
Detailed Description
In order to make the objects, technical solutions and advantages of the present application more apparent, the present application will be further described in detail with reference to the accompanying drawings and examples. It should be understood that the specific embodiments described herein are for purposes of illustration only and are not intended to limit the present application.
The invention provides an on-orbit target approaching guiding method based on an attitude-orbit coupling control strategy, which reduces the influence of an attitude angle on the acceleration of instructions in a guidance law in the whole end guiding process, and simultaneously reduces the problem of attitude-orbit correction coupling by utilizing a time-sharing multiplexing theory. Referring to fig. 1, the specific steps are as follows:
step 1: establishing a relative motion model of a spacecraft and a target
In simulation verification of an on-orbit target approach guidance method based on a pose-orbit coupling control strategy, measurement information needs to be acquired as input quantity of the guidance method, so that a relative motion model of a spacecraft and a target needs to be established to acquire the measurement information. The measurement information comprises an azimuth angle and a high-low angle of a relative sight line of the spacecraft and the target, a yaw angle, a pitch angle and a roll angle for describing the posture of the spacecraft, and the speed of the spacecraft, the azimuth angle and the high-low angle of the speed, wherein the initial value also comprises an initial distance between the spacecraft and the target.
And calculating the approaching speed and the relative distance between the spacecraft and the target at each moment in the control period according to the measured value, the relative motion model and the initial relative distance. In order to study the design method of the tail end guidance law of the spacecraft in the three-dimensional space, the relative motion relation of the spacecraft M and the target T is split into a horizontal plane and a vertical plane under an inertial coordinate system to study the horizontal plane and the vertical plane respectively, wherein the horizontal plane XOY is the XOY plane of the inertial coordinate system projected with the relative position relation of the spacecraft and the target, the vertical plane yoz is the same vertical plane where the spacecraft and the target are always located, and the geometric relation of the spacecraft and the target is shown in fig. 5.
Where oy is the projection of the line of sight MT on the horizontal plane XOY, α is the line of sight azimuth, and β is the line of sight pitch angle. The position vector of the spacecraft M and the target T relative to the earth center in the inertial coordinate system is R m 、R t . The target moves under the action of gravity without generating maneuver, and the passive external force applied to the spacecraft comprises gravity and cold air jet thrust. According to newton kinematics, the relative motion model between the two can be expressed as:
wherein A is t For target acceleration caused by earth's gravity, A m G= 6.67259 ×10 for spacecraft acceleration due to gravitational attraction and cold air jet thrust -11 N·m 2 /kg 2 ，M d ＝5.965×10 24 kg is the gravitational constant and the earth mass respectively, A control Acceleration is commanded for control acting on the spacecraft. From fig. 5, the relative position vector R and the absolute position vector R of the spacecraft and the target can be obtained t 、R m Is R=R t -R m 。
The azimuth angle and the high-low angle of the spacecraft speed and the azimuth angle and the high-low angle of the approaching speed of the spacecraft and the target can be obtained according to the motion relation of the spacecraft and the target in the three-dimensional space, and the azimuth angle and the high-low angle of the approaching speed of the spacecraft and the target are as follows:
wherein alpha is vm 、β vm Azimuth angle and pitch angle of spacecraft velocity vector respectively, alpha v 、β v Respectively, an azimuth angle and a pitch angle which approach the velocity vector, R mx 、R my 、R mz Respectively absolute position vectors R m Projection on x, y, z axis, R x 、R y 、R z The projections of the relative position vector R on the x, y, z axes, respectively.
Step 2: establishing a spacecraft attitude dynamics model
According to the structural layout of the spacecraft air nozzle symmetrical about the centroid and the rotation dynamics equation of the longitudinal axis around the centroid, a spacecraft rigid body attitude angle dynamic model is established, and the yaw angular acceleration of the spacecraft is obtainedIs +.>Regarding the cold air spraying force F of the horizontal plane and the vertical plane under the inertial coordinate system a 、F b Is a function of the yaw rate w of the spacecraft a Angular velocity w from pitch angle b A nonlinear function variation relationship with time t.
The structure layout of the spacecraft with the Leng Qipen force symmetrical about the mass center is shown in fig. 1, and two pairs of cold air nozzles at one side of the mass center of the spacecraft respectively control yaw and pitch channels of the spacecraft on two planes of a horizontal plane and a vertical plane under an inertial coordinate system. To describe the change condition of the attitude angle of the space spacecraft in a quasi-real time, the yaw angular acceleration of the space spacecraft can be established according to the motion characteristic of the longitudinal axis of the space spacecraft in the space and the cold air spraying force structural layoutIs +.>The change relation between the resultant force of the horizontal plane and the resultant force of the vertical plane in the inertial coordinate system is related.
Regarding a spacecraft running in space as a cylindrical rigid body, according to the rigid body fixed-axis rotation law, when forces which are symmetrical about the centroid of the spacecraft and opposite in direction act on the spacecraft in the direction perpendicular to the longitudinal axis of the spacecraft, the forces are projected under the horizontal plane in an inertial coordinate system to form resultant force F a Projection forming resultant force F under vertical plane b Generating a resultant moment M acting on the spacecraft a M and M b Which in turn produces angular acceleration of rotation about the centroidIs->Yaw rate w a Angular velocity w from pitch angle b Slope with time t +.>Is->Is a linear variation relationship of (2); when the force applied to the spacecraft is 0, the attitude motion state of the spacecraft operated in the vacuum environment is kept unchanged, namely the yaw rate w of the spacecraft a Angular velocity w from pitch angle b Is constant and does not change with time t. When the force projected in the direction perpendicular to the longitudinal axis of the spacecraft is the resultant force of-F under the horizontal plane and the vertical plane in the inertial coordinate system a and-F b Generating an angular acceleration of rotation about the centroid>Is->Yaw rate w a Angular velocity w from pitch angle b Slope with time t +.>Is->The linear change relation of (2) can be specifically expressed as follows:
wherein J is the moment of inertia of the spacecraft around the centroid, m 0 Is the total mass of the spacecraft, R is the cross-section diameter of the spacecraft, l is the total length of the spacecraft,the angular acceleration of the spacecraft rotating around the centroid is F, the resultant force of the spraying forces of the spacecraft, which are symmetrical about the centroid and have opposite front and rear directions, r is the distance between the spraying force of the spacecraft and the centroid, and M is the resultant moment generated by the resultant force F.
The yaw acceleration and pitch acceleration of a spacecraft as a function of the resultant force produced by the controlling spacecraft can be expressed as:
wherein F is a 、F b Are constant positive integers greater than zero.
Yaw rate w of spacecraft a Angular velocity w from pitch angle b The nonlinear functional change relation with time t can be expressed as:
wherein w is a 、w a The yaw angular velocity and the pitch angular velocity, w, of the spacecraft a0 、w b0 The initial values of the yaw angular velocity and the pitch angular velocity of the spacecraft are obtained.
Step 3: design of attitude and orbit correction control strategy for attitude and orbit correction coupling problem
According to yaw angular acceleration of spacecraftIs +.>Leng Qipen force F with respect to horizontal and vertical planes in inertial frame a 、F b Is a function of the yaw rate w of the spacecraft a Angular velocity w from pitch angle b The control strategy is designed in a nonlinear function change relation with time t. Because the cold air spraying force of the spacecraft is symmetrical about the centroid and the gesture correction and track guidance of the spacecraft generate a coupling problem, the gesture correction and the track guidance of the spacecraft cannot be synchronously performed, a control strategy is designed according to a nonlinear function of the angular velocity change relation about the longitudinal axis of the spacecraft, so that the gesture correction can be realized in the whole tail end guiding process of the spacecraft, and the requirements of ensuring the target in the view field range and the track guidance can be realized, and ensuring the miss distance can be met.
Based on the time-sharing multiplexing theory, when the spacecraft yaw angular velocity w a Or pitch angle angular velocity w b When the position is 0, controlling the spacecraft to conduct track guidance only, and defining the track guidance as a track correction stage 0; when the absolute value of the deviation between the relative sight angle of the spacecraft and the target and the attitude angle of the spacecraft exceeds the central threshold gamma of the field of view a 、γ b When the spacecraft is controlled to perform attitude correction only, the yaw rate w a Angular velocity w from pitch angle b Slope with time t ofIs->Defining a linear change relation of the attitude correction stage 1; setting w according to spacecraft attitude angle correction accuracy a And w is equal to b W of maximum value of (2) amax 、w bmax When the yaw rate w of the spacecraft is generated a Angular velocity w from pitch angle b When the yaw angle is nonzero and reaches the maximum value, the spacecraft is controlled to conduct track guidance only, and at the moment, the yaw angle theta or the pitch angle of the spacecraft is controlled to be +>At angular velocity w amax 、w bmax Continuously correcting, namely correcting the track according to a designed guidance law, and defining the track as a track correction stage 2 in a track composite correction stage; when the absolute value of the deviation between the relative sight angle of the spacecraft and the target and the attitude angle of the spacecraft enters the central threshold gamma of the field of view a 、γ b In the inner time, controlling the yaw rate w of the spacecraft a Angular velocity w from pitch angle b Slope with time t +.>Is->Is corrected for w by linear change relation of (2) a Or w b And 0, wherein the spacecraft is in a posture correction force recoil stage, which is defined as a posture correction force recoil stage 3.
As shown in fig. 2, a flowchart of the primary attitude correction control strategy is that the change relation between the yaw angular velocity and the pitch angular velocity of the spacecraft attitude with respect to time t can be represented by the following formula:
wherein, due to the constraint of the control period T, the spacecraft can only send out instructions at the kT moment, so the critical moment of each control stage meets T 1 、t 2 、t 3 、t 4 、t' 1 、t' 2 、t' 3 、t' 4 =kt (k=0, 1,2, 3.). In order to ensure that the yaw angle and the pitch angle are corrected to 0 in the attitude correction force recoil stage 3, t is required to be satisfied 2 -t 1 ＝t 4 -t 3 ,t' 2 -t' 1 ＝t' 4 -t' 3 . The curve of the change of the once correction process of the yaw angular velocity and the pitch angular velocity of the spacecraft with time t is shown in fig. 3.
Track correction stage (stage 0): when the spacecraft does not deflect, the cold air valves on two sides of the mass center are simultaneously opened by the spacecraft, and the thrust acts on the mass center to perform pure tail end guiding movement;
posture correction stage (stage 1): when the attitude angle of the spacecraft is not in the required range, adjusting a cold air valve of the spacecraft, opening a valve at one side of a centroid to generate a rotational moment of the spacecraft body, so that the rotational angle speed of the spacecraft in t 1-t 2 is in a direct proportion relation with time, and the running state of the spacecraft in the stage 1 is the attitude correction stage;
posture rail correction stage (stage 2): taking into consideration that the rotation angular velocity of the spacecraft increases to an upper limit, opening cold air valves at two sides of the mass center, and enabling the spacecraft to self at the angular velocity w max The spacecraft performs pure tail end guiding movement while continuously rotating, and the spacecraft is in a stage 2 at this time, namely a pose-orbit composite correction stage;
posture correction force recoil phase (phase 3): considering that the spacecraft operates in a vacuum space, the angular velocity w required to be generated for the spacecraft max And (3) counteracting the generated rotation angular velocity by using cold air recoil force, opening a cold air valve on one side of the mass center at the moment, and enabling the cold air valve to be opposite to the cold air spraying force opened in the stage 1, wherein the spacecraft is in the stage 3, namely the posture correction force recoil stage.
In order to ensure that the target is always positioned at the center of the field of view in the running process of the spacecraft and the effect of the pure tail end guiding correction force is maximum, the correction attitude angle is considered to be preferentially considered when the attitude angle deviates from the constraint range, and the change of the attitude angle after correction and recovery is small in the tail end guiding section due to the short relative distance between the spacecraft and the target, so that the change influence can be ignored.
Center threshold gamma of field of view a 、γ b Is of the size of (d) and w amax 、w bmax The size of (2) may be determined by the following formula:
wherein delta theta,The yaw angle and pitch angle of the spacecraft are changed in the attitude correction stage 0 of the control strategy. Delta epsilon a 、Δε b The deviation of the yaw angle and the pitch angle of the longitudinal axis of the spacecraft and the azimuth angle and the pitch angle of the relative sight direction.
Step 4: introducing auxiliary variable related to attitude angle of spacecraft to improve generalized proportional guidance law
Introducing an auxiliary variable epsilon with respect to attitude angle of spacecraft a 、ε b Correcting the command acceleration a of the plane of the yaw angle of the spacecraft a Commanded acceleration a of plane in which pitch angle is located b 。
For a track correction part in a control strategy, since the instruction acceleration direction calculated according to the generalized proportion guidance law is perpendicular to the sight line direction, the acting force of the control spacecraft is perpendicular to the longitudinal axis direction, the longitudinal axis has angle deviation with the sight line, and an auxiliary variable epsilon related to the attitude angle of the spacecraft is introduced to avoid the influence of the action force on the off-target amount a 、ε b Correcting the command acceleration a of the plane of the yaw angle of the spacecraft a Commanded acceleration a of plane in which pitch angle is located b . The schematic diagram of the relative sight line of the spacecraft and the target and the generalized proportional guiding command acceleration of the spacecraft and the longitudinal axis of the spacecraft is shown in fig. 4.
Firstly, according to measurement information obtained by a relative motion model of a spacecraft and a target, calculating the approaching speed and the relative line-of-sight azimuth angle change rate of a horizontal plane of the spacecraft under an inertial coordinate system, and the approaching speed and the relative line-of-sight pitch angle change rate of a vertical plane under the inertial coordinate system, wherein the approaching speed and the relative line-of-sight pitch angle change rate are shown in the following formula:
wherein R is x 、R y 、R z The projections of the relative position vector R of the spacecraft and the target on the x, y and z axes are respectively shown. Alpha is the relative azimuth angle of the line of sight of the spacecraft and the target.The approach speeds of the spacecraft and the target are projected on a horizontal plane and a vertical plane in an inertial coordinate system respectively. The rate of change of the relative attitude angle α and pitch angle β of the spacecraft to the target can be expressed as follows:
the instruction acceleration which can be obtained in the yaw plane and the pitch plane according to the generalized proportional guidance law is as follows:
wherein N is 1 、N 2 The proportional guide coefficients of the yaw plane and the pitch plane are respectively. The commanded acceleration of the yaw plane is located at a horizontal plane under the inertial coordinate system and perpendicular to the relative line of sight direction, and the commanded acceleration of the pitch plane is located at a vertical plane under the inertial coordinate system and perpendicular to the relative line of sight direction. In order to consider that the command acceleration output by the actual spacecraft is perpendicular to the longitudinal axis direction, an auxiliary variable related to the attitude angle of the spacecraft is introduced to compensate errors.
As can be seen from fig. 4, the output command acceleration for improving the generalized proportional guidance law is:
step 5: PWM pulse ignition strategy design according to improved generalized proportional pilot law
Because the control period exists and the cold air spraying force is constant, the output command acceleration designed by the guidance law cannot be used as the spacecraft control output command acceleration, so that a PWM pulse ignition strategy is required to be designed, and the spacecraft running in space gives a thrust F a F (F) b Is a constant value, and because the instruction acceleration calculated by the guidance law is a time-varying function related to time t, and the trigger of the spacecraft instruction acceleration has unavoidable control period constraint, the PWM pulse ignition strategy can be designed according to the Schmitt trigger logic.
The acceleration w can be calculated from the constant thrust m As a switching interval threshold:
at the beginning of a control period Tinitial time, judging whether the command acceleration is positioned in a switch interval, controlling the output of the command acceleration, wherein the command acceleration is positioned in the switch interval, and the output command acceleration is w m And estimating the duty ratio of the output command acceleration in a control period T through a relative motion model of the spacecraft and the target, wherein the command acceleration is positioned in a related interval, and the output command acceleration is 0 in the control period T. A PWM pulse function of the commanded acceleration with respect to time t is established. The final output pulse command acceleration can be expressed as:
where k=0, 1,2,3.
The invention also provides an on-orbit target approach guidance system based on the gesture-orbit coupling control strategy, and the on-orbit target approach guidance based on the gesture-orbit coupling control strategy is realized based on the on-orbit target approach guidance method.
A computer device comprises a memory, a processor and a computer program stored on the memory and capable of running on the processor, wherein when the processor executes the computer program, on the basis of the on-orbit target approach guidance method, on-orbit target approach guidance based on a gesture-orbit coupling control strategy is realized.
A computer readable storage medium having stored thereon a computer program which, when executed by a processor, implements on-orbit target approach guidance based on an on-orbit target approach guidance method described.
In summary, the given thrust F of the spacecraft which operates in space can be known according to the spacecraft attitude dynamics model a F (F) b The method is a constant value, because the instruction acceleration calculated by the improved generalized proportional guidance law is a time-varying function related to time t, and unavoidable control period constraint exists in the triggering of the instruction acceleration of the spacecraft, the improved pulse generalized proportional guidance law can better act on the horizontal plane and the vertical plane under the inertial coordinate system of the spacecraft, and finally the track correction part in the control strategy achieves better off-target accuracy by utilizing the improved pulse guidance method.
Examples
To verify the effectiveness of the inventive protocol, the following experiments were performed.
1. Simulation conditions
And the spacecraft enters an end guiding stage when the distance between the spacecraft and the target reaches 5 km, the power mode of the spacecraft is changed into cold air correction by taking the centroid point of the spacecraft as a symmetry center, and the cold air correction is concentrated in the horizontal and vertical directions under an inertial coordinate system. Target orbit height 36000.22m, spacecraft orbit height 36003.25m. The initial position and speed of the tail end guide of the spacecraft are state information of the spacecraft at a distance of 5 km from the target, and the initial attitude angle and the initial speed angle of the spacecraft body are consistent.
Assuming that the initial position and velocity of the target at the start of the end guide are:
the initial position and the initial speed of the spacecraft are as follows:
initial yaw angle and pitch angle of spacecraft are theta 0 ＝224.2°,The roll angle is always 0, the initial relative distance between the spacecraft and the target is 4990m, and the azimuth angle and the altitude angle of the relative sight of the spacecraft and the target are alpha 0 ＝189.2°,β 0 The measurement angle error is 0.1 degrees, the initial relative speed and direction of the spacecraft and the target can be expressed as follows:
control period T:250ms; leng Qipen force F a 、F b :100mN; leng Qipen force attenuation error: -30%; the terminal guidance law proportionality coefficient is set to be (6, 6) in the horizontal plane and the vertical plane; horizontal plane and vertical plane attitude correction period time: (2 t,2 t); the approach speed is 103.88m/s; threshold value gamma of central angle of field of view a 、γ b Is + -2 deg..
2. Simulation content and result analysis
As shown in fig. 6, which is a three-dimensional track diagram of a spacecraft and a target, the process of guiding the spacecraft to the target according to a set guiding law can be obviously observed; FIG. 7 shows the relative distance between the spacecraft and the target over time, and it can be found that the off-target amount of the spacecraft can reach high accuracy within 0.554m in the algorithm of the end guidance law designed by the invention; FIGS. 8 and 9 show the magnitude of the spraying force of a pair of cold air spouts in each control period T in the horizontal direction and the vertical direction; FIGS. 10 and 11 show the angular difference between the attitude angle and the line of sight angle of the spacecraft, and the research shows that the angular difference can be corrected to be within + -2 DEG of the center of the field of view, so that the target is always positioned in the center of the field of view and is not lost, and the spacecraft can be ensured to be aligned with the target in the head direction; fig. 12 and fig. 13 show the time-dependent relationship between the rotational angular velocity of the attitude angle of the spacecraft, and the coordinated transition of the working state of the spacecraft in the end guidance law designed by the invention can be found by observation.
In summary, the on-orbit target approach guidance method based on the pose-orbit coupling control strategy provided by the invention is more suitable for actual engineering application of a spacecraft approaching a non-cooperative target in space.
The technical features of the above embodiments may be arbitrarily combined, and all possible combinations of the technical features in the above embodiments are not described for brevity of description, however, as long as there is no contradiction between the combinations of the technical features, they should be considered as the scope of the description.
The above examples only represent a few embodiments of the present application, which are described in more detail and are not to be construed as limiting the scope of the present application. It should be noted that it would be apparent to those skilled in the art that various modifications and improvements could be made without departing from the spirit of the present application, which would be within the scope of the present application. Accordingly, the scope of protection of the present application shall be subject to the appended claims.
Claims (8)
1. An on-orbit target approach guidance method based on an attitude-orbit coupling control strategy is characterized by comprising the following steps:
step 1: establishing a spacecraft attitude dynamics model
According to aerospaceThe structural layout of the air jet of the spacecraft about the symmetry of the centroid and the rotation dynamics equation of the longitudinal axis about the centroid are established, and a rigid body attitude angle dynamic model of the spacecraft is established to obtain yaw angular acceleration of the spacecraftIs +.>Regarding the cold air spraying force F of the horizontal plane and the vertical plane under the inertial coordinate system a 、F b Is a function of the yaw rate w of the spacecraft a Angular velocity w from pitch angle b A nonlinear function change relation with time t;
step 2: construction of attitude and orbit correction control strategy aiming at attitude and orbit correction coupling problem
According to yaw angular acceleration of spacecraftIs +.>Leng Qipen force F with respect to horizontal and vertical planes in inertial frame a 、F b Is a function of the yaw rate w of the spacecraft a Angular velocity w from pitch angle b Constructing a pose rail correction control strategy according to the nonlinear function change relation of time t, and ensuring that the relative sight angle is in the center of the view field of the spacecraft;
step 3: introducing auxiliary variable related to attitude angle of spacecraft to improve generalized proportional guidance law
Introducing an auxiliary variable epsilon with respect to attitude angle of spacecraft a 、ε b Correcting instruction acceleration a of plane where yaw angle of spacecraft in generalized proportional guidance law is located a Commanded acceleration a of plane in which pitch angle is located b To reduce the error of the instruction acceleration caused by the existence of the attitude angle;
step 4: PWM pulse ignition strategy design according to improved generalized proportional pilot law
According to the plane instruction acceleration a of the yaw angle of the spacecraft a The plane command acceleration a of pitch angle b Regarding the continuous time-varying function change relation of time t, combining with Schmitt trigger logic to construct a PWM pulse ignition strategy, so that a spacecraft carries out track correction according to an improved pulse generalized proportion guiding method to achieve high-precision off-target quantity;
step 1, establishing a spacecraft attitude dynamics model, wherein the specific method comprises the following steps:
according to the structural layout of the spacecraft air nozzle symmetrical about the centroid and the rotation dynamics equation of the longitudinal axis around the centroid, a spacecraft rigid body attitude angle dynamics model is established, and the model is shown in the following formula:
wherein J is the moment of inertia of the spacecraft around the centroid, m 0 Is the total mass of the spacecraft, R is the cross-section diameter of the spacecraft, l is the total length of the spacecraft,the angular acceleration of the spacecraft rotating around the centroid is F, the resultant force of the spraying forces of the spacecraft, which are symmetrical about the centroid and have opposite front and rear directions, r is the distance between the spraying force of the spacecraft and the centroid, and M is the resultant moment generated by the resultant force F;
yaw angular acceleration of spacecraftIs +.>Regarding the cold air spraying force F of the horizontal plane and the vertical plane under the inertial coordinate system a 、F b Is shown in the following formula:
wherein F is a 、F b Are constant positive integers greater than zero;
when the resultant force of the force projected in the direction perpendicular to the longitudinal axis of the spacecraft under the horizontal plane and the vertical plane in the inertial coordinate system is F a And F is equal to b When generating a resultant moment M acting on the spacecraft a M and M b Which in turn produces angular acceleration of rotation about the centroidIs->At this time yaw rate w a Angular velocity w from pitch angle b Slope with time t +.>Is->Is a linear variation relationship of (2); when the force applied to the spacecraft is 0, the attitude motion state of the spacecraft operated in the vacuum environment is kept unchanged, namely the yaw rate w of the spacecraft a Angular velocity w from pitch angle b Is constant and does not change with time t; when the force projected in the direction perpendicular to the longitudinal axis of the spacecraft is the resultant force of-F under the horizontal plane and the vertical plane in the inertial coordinate system a and-F b When an angular acceleration of rotation about the centroid is generated +.>Is->Yaw rate w a Angular velocity w from pitch angle b Slope with time t +.>Is->Is a linear change of the yaw rate w of the spacecraft a Angular velocity w from pitch angle b The nonlinear functional change with time t is shown in the following equation:
wherein w is a 、w a The yaw angular velocity and the pitch angular velocity, w, of the spacecraft a0 、w b0 The initial values of the yaw angular velocity and the pitch angular velocity of the spacecraft are obtained.
2. The on-orbit target approach guidance method based on the pose-orbit coupling control strategy according to claim 1, wherein step 2, a pose-orbit correction control strategy is constructed for the pose-orbit correction coupling problem, and the specific method is as follows:
based on the time-sharing multiplexing theory, when the spacecraft yaw angular velocity w a Or pitch angle angular velocity w b When the position is 0, controlling the spacecraft to conduct track guidance only, and defining the track guidance as a track correction stage 0; when the absolute value of the deviation between the relative sight angle of the spacecraft and the target and the attitude angle of the spacecraft exceeds the central threshold gamma of the field of view a 、γ b When the spacecraft is controlled to perform attitude correction, the generated acting force is F a 、F b So that the yaw rate w a Angular velocity w from pitch angle b Slope with time t ofIs->Defining a linear change relation of the attitude correction stage 1; when the yaw angular velocity w of the spacecraft a Angular velocity w from pitch angle b When the yaw angle is nonzero and reaches the maximum value, the spacecraft is controlled to conduct track guidance only, and at the moment, the yaw angle theta or the pitch angle of the spacecraft is controlled to be +>At maximum angular velocity w amax 、w bmax Continuously correcting, namely correcting the track according to the set guidance law, and defining the track as a track correction stage 2 in a track composite correction stage; when the absolute value of the deviation between the relative sight angle of the spacecraft and the target and the attitude angle of the spacecraft enters the central threshold gamma of the field of view a 、γ b When in use, the spacecraft is controlled to generate acting force of-F a 、-F b Yaw rate w a Angular velocity w from pitch angle b Slope with time t +.>Is->Is corrected for w by linear change relation of (2) a Or w b For 0, the spacecraft is in a posture correction force recoil section, which is defined as a posture correction force recoil section 3, and a posture rail correction control strategy is specifically shown as follows:
because of the constraint of the control period T, the spacecraft can only send out instructions at the kT moment, so the critical moment of each control stage meets T 1 、t 2 、t 3 、t 4 、t 1 '、t' 2 、t 3 '、t' 4 Let kT (k=0, 1,2,3,) and to ensure that the attitude correction force recoil phase 3 ensures that the yaw and pitch angles are corrected to 0, t is satisfied 2 -t 1 ＝t 4 -t 3 ,t' 2 -t 1 '＝t' 4 -t' 3 。
3. The on-orbit target approach guidance method based on the pose-orbit coupling control strategy according to claim 2, wherein the field-of-view center threshold γ a 、γ b Is of the size of (d) and w amax 、w bmax The size of (2) is determined by the following formula:
wherein delta theta,For the yaw angle and pitch angle of the spacecraft which are changed in the attitude correction stage 0 of the control strategy, delta epsilon a 、Δε b The deviation of the yaw angle and the pitch angle of the longitudinal axis of the spacecraft and the azimuth angle and the pitch angle of the relative sight direction.
4. The on-orbit target approach guidance method based on the attitude and orbit coupling control strategy according to claim 3, wherein the step 3 is to introduce auxiliary variables related to the attitude angle of the spacecraft to improve the generalized proportional guidance law, and the specific method is as follows:
firstly, according to a relative motion model of a spacecraft and a target, calculating the approaching speed and the relative line-of-sight azimuth angle change rate of a horizontal plane of the spacecraft under an inertial coordinate system, and the approaching speed and the relative line-of-sight pitch angle change rate of a vertical plane under the inertial coordinate system, wherein the approaching speed and the relative line-of-sight pitch angle change rate are shown in the following formula:
wherein R is x 、R y 、R z Respectively the projections of relative position vectors R of the spacecraft and the target on x, y and z axes, alpha is the relative line-of-sight azimuth angle of the spacecraft and the target,the approach speeds of the spacecraft and the target are projected on a horizontal plane and a vertical plane under an inertial coordinate system respectively;
the change rate of the relative attitude angle alpha and pitch angle beta of the spacecraft and the target is shown as follows:
obtaining the instruction acceleration a in the yaw plane and the pitch plane according to the generalized proportional guidance law ac 、a ac The method comprises the following steps:
wherein N is 1 、N 2 The proportional guide coefficients are respectively a yaw plane and a pitch plane, the command acceleration of the yaw plane is positioned on a horizontal plane under an inertial coordinate system and is perpendicular to the relative sight line direction, and the command acceleration of the pitch plane is positioned on a vertical plane under the inertial coordinate system and is perpendicular to the relative sight line direction;
in order to consider the direction of the command acceleration of the actual spacecraft output perpendicular to the longitudinal axis, an auxiliary variable epsilon related to the attitude angle of the spacecraft is introduced a 、ε b Compensating errors;
the improved spacecraft output instruction acceleration is as follows:
5. the on-orbit target approach guidance method based on the gesture-orbit coupling control strategy according to claim 4, wherein the step 4 is to design a PWM pulse ignition strategy according to an improved generalized proportional guidance law, and the specific method is as follows:
first, the generated acceleration w is calculated from the constant thrust m As a switching interval threshold value:
then establish the commanded acceleration a' a 、a' b PWM pulse function for time t:
at the beginning of a control period Tinitial time, judging whether the command acceleration is positioned in a switch interval, controlling the output of the command acceleration, wherein the command acceleration is positioned in the switch interval, and the output command acceleration is w m Estimating the duty ratio of the output command acceleration in a control period T through a relative motion model of the spacecraft and the target, wherein the command acceleration is positioned in an off interval, and the output command acceleration is 0 in the control period T, and is shown in the following formula:
where k=0, 1,2,3.
6. An on-orbit target approach guidance system based on an attitude-orbit coupling control strategy is characterized in that on-orbit target approach guidance based on the attitude-orbit coupling control strategy is realized based on the on-orbit target approach guidance method of any one of claims 1 to 5.
7. A computer device comprising a memory, a processor and a computer program stored on the memory and executable on the processor, the processor implementing on-orbit target approach guidance based on a gesture-orbit coupling control strategy based on the on-orbit target approach guidance method of any one of claims 1-5 when executing the computer program.
8. A computer readable storage medium having stored thereon a computer program which, when executed by a processor, implements on-orbit target approach guidance based on a gesture-orbit coupling control strategy based on the on-orbit target approach guidance method of any one of claims 1-5.
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