A rigid spacecraft attitude tracking method and related products

Abstract

The present invention discloses a rigid spacecraft attitude tracking method and related products, belonging to the field of spacecraft control technology. The rigid spacecraft attitude tracking method provided by the present invention, by introducing a preset time switching function as a time-varying gain, adjusts the attitude control torque and the adaptive law of the uncertain inertia identification process, ensures that the spacecraft completes attitude tracking within a preset time and maintains tracking after a preset adjustment time, thereby improving the accuracy and reliability of control; when the inertia matrix parameters are unknown, the spacecraft can complete precise attitude tracking control within a preset time, ensure that the torque input remains continuous and bounded, avoid chattering, and maintain zero tracking error after a preset adjustment time, thus having high engineering application value; this method does not require a complex parameter adjustment process, and the switching function structure is independent of the initial value of the system state, greatly improving the compatibility and effectiveness of the control method, and can adapt to different spacecraft and working conditions.

Description

Technical Field

[0001] The present invention relates to the field of spacecraft control technology, and in particular to a rigid spacecraft attitude tracking method and related products. Background Art

[0002] With the rapid development of space exploration technology, spacecraft are increasingly being used in fields such as space telescope observation, remote sensing satellite data collection, and deep space exploration. These missions place increasing demands on the complexity and performance of spacecraft, particularly in attitude maneuvering control. Attitude maneuvers require spacecraft to accurately and stably transition from their current attitude to a desired attitude, while also dynamically tracking the target.

[0003] In existing spacecraft attitude tracking control systems, the design of attitude control algorithms is highly dependent on the spacecraft's inertia matrix, a key factor in determining control torque. However, during a spacecraft's mission, its structure may change, such as the release of cargo, the installation or removal of equipment, and the replacement of components. These changes alter the spacecraft's mass distribution, which in turn affects its inertia matrix. Furthermore, spacecraft face significant temperature fluctuations in the space environment, particularly the temperature difference between sunlit and shaded areas. These temperature fluctuations cause the spacecraft's materials to expand or contract, further altering its mass distribution and inertia matrix.

[0004] Changes in the inertia matrix pose a significant challenge to the stability and accuracy of attitude control systems. If the control system cannot accurately identify changes in the inertia matrix, errors will occur in the spacecraft's angular velocity and attitude control. These errors can gradually accumulate, ultimately leading to a degradation of control performance or even mission failure. Therefore, existing attitude control systems have significant shortcomings in coping with inertia matrix changes.

[0005] In order to solve the above problems, it is necessary to develop a method that can ensure that the spacecraft completes precise attitude tracking control within a preset time when the inertia matrix parameters are unknown. Summary of the Invention

[0006] The purpose of the present invention is to provide a rigid spacecraft attitude tracking method and related products to overcome the problem of inaccurate attitude tracking caused by changes in the inertia matrix in the prior art.

[0007] The present invention solves the above technical problems through the following technical solutions:

[0008] A rigid spacecraft attitude tracking method comprises the following steps:

[0009] Step 1: Based on the attitude kinematics and dynamics equations of the rigid spacecraft body, the attitude kinematics and dynamics equations of the target rigid spacecraft are constructed to obtain the attitude quaternion error of the rigid spacecraft body relative to the target rigid spacecraft and the angular velocity error of the rigid spacecraft body coordinate system relative to the target rigid spacecraft coordinate system;

[0010] Step 2: Construct a switching function with respect to a preset time, establish a virtual controller based on the attitude quaternion error of the rigid spacecraft body relative to the target rigid spacecraft, and calculate the derivative of the virtual controller;

[0011] Step 3: Based on the attitude quaternion error of the rigid spacecraft body relative to the target rigid spacecraft and the angular velocity error of the rigid spacecraft body coordinate system relative to the target rigid spacecraft coordinate system, the three-axis control torque required to complete angular velocity tracking within a preset time is calculated in combination with the virtual controller and its derivatives;

[0012] Step 4: Perform matrix dimension deformation on the dynamic equation of the rigid spacecraft body to obtain the differential deformation equation; based on the three-axis control torque, perform time-invariant gain filtering on the variables in the differential deformation equation to obtain the first algebraic equation; perform time-invariant gain filtering on the variables in the first algebraic equation to obtain the second algebraic equation; perform time-varying gain filtering on the second algebraic equation to obtain the third algebraic equation;

[0013] Step 5: Based on the third algebraic equation, establish an adaptive law for the inertia matrix parameters of the rigid spacecraft body within a preset time, which is used to realize the inertia matrix parameter identification and realize the attitude tracking of the rigid spacecraft body.

[0014] A further improvement of the present invention is that, in step 1, the attitude kinematic equation of the rigid spacecraft body is:

[0015]

[0016] Among them, q(t) is the attitude quaternion of the rigid spacecraft body coordinate system relative to the inertial reference coordinate system at time t, q0(t) is the scalar part of q(t), and q v (t) is the vector part of q(t), is the first-order derivative of q(t) with respect to time, ω(t) is the angular velocity vector of the spacecraft body coordinate system relative to the inertial reference coordinate system at time t, and I3 is the unit matrix;

[0017] The attitude dynamics equation of the rigid spacecraft body is:

[0018]

[0019] in, is the first-order derivative of ω(t) with respect to time, J is the unknown moment of inertia matrix of the rigid spacecraft body, which is a symmetric matrix; u(t) is the three-axis control torque applied to the rigid spacecraft body;

[0020] The attitude kinematic equation of the target rigid spacecraft is:

[0021]

[0022] Among them, q d (t) is the attitude quaternion of the target rigid spacecraft coordinate system relative to the rigid spacecraft body coordinate system at time t, where q d0 (t) is q d The scalar part of (t), q dv (t) is q d The vector part of (t), q d (t) The first derivative with respect to time, ω d (t) is the angular velocity vector in the target rigid spacecraft coordinate system relative to the rigid spacecraft body coordinate system at time t;

[0023] The attitude dynamics equation of the target rigid spacecraft is:

[0024]

[0025] in, ω d (t) the first derivative with respect to time;

[0026] Attitude quaternion error of the rigid spacecraft body relative to the target rigid spacecraft for:

[0027]

[0028] Angular velocity error of the rigid spacecraft body coordinate system relative to the target rigid spacecraft coordinate system for:

[0029]

[0030] in, is the transformation matrix from the target rigid spacecraft coordinate system to the rigid spacecraft body coordinate system at time t.

[0031] A further improvement of the present invention is that the switching function K(t) is:

[0032]

[0033] Among them, t f is the preset time.

[0034] A further improvement of the present invention is that the virtual controller α(t) is:

[0035]

[0036] Among them, c1 is a normal constant that can be designed; for The vector part of

[0037] Derivatives of virtual controllers for:

[0038]

[0039] in, for The scalar part of is the first-order derivative of the switching function K(t) at time t, specifically:

[0040]

[0041] A further improvement of the present invention is that the three-axis control torque u(t) for completing angular velocity tracking within a preset time is calculated as follows:

[0042]

[0043] Among them, c2 is a normal number that can be designed. is the unknown inertia estimation matrix, Specifically:

[0044]

[0045] in, of They are the unknown inertia matrix J at time t. 11 , J 22 , J 33 , J 23 , J 13 , J 12 estimated value.

[0046] A further improvement of the present invention is that, in step 4, [J 11 J 22 J 33 J 23 J 13 J 12 ] T is the intermediate matrix θ, and the differential transformation equation is:

[0047]

[0048] Among them, W(t) is the intermediate parameter, specifically:

[0049] W(t)=-ω × (t)L(ω(t))

[0050] Where, L(ω(t)) is the matrix about ω(t);

[0051] The variables in the differential transformation equation are filtered with time-invariant gain, specifically:

[0052]

[0053] Among them, k1 is a normal constant that can be designed, ω f1 (t), W f1 (t) and u f1 (t) are the output signals of the filter after the time-invariant gain filtering of the variables in the differential transformation equation at time t, and are u at time t f1 (t),ω f1 (t) and W f1 (t) the first derivative with respect to time;

[0054] Set variable The first algebraic equation is:

[0055]

[0056] For variables in first algebraic equations and u f1 (t) Perform filtering with time-invariant gain, specifically:

[0057]

[0058] Where k2 is a normal number that can be designed. and u f2 (t) are all variables in the first algebraic equation at time t and u f1 (t) performing time-invariant gain filtering on the output signal of the filter; and They are u f2 (t) The first derivative with respect to time, specifically:

[0059] The equation is:

[0060]

[0061] definition is time t The adjoint matrix of is time t The value of the determinant is:

[0062]

[0063] The second algebraic equation is obtained:

[0064] Y(t)=Δ(t)θ

[0065] The second algebraic equation is filtered with a time-varying gain, where the time-varying gain is the square of Δ(t), i.e., Δ 2 (t), specifically:

[0066]

[0067] Among them, Δ f (t) is the output signal after the second algebraic equation is filtered by time-varying gain, and Δ f (t) is the adaptive gain of the spacecraft inertia matrix parameter identification algorithm, Y f (t) is the output signal of the filter after time-varying gain filtering of the second algebraic equation; and Δ is the time t f (t), Y f The first-order derivative of (t) with respect to time gives the third algebraic equation:

[0068] Y f (t) = Δ f (t)θ.

[0069] A further improvement of the present invention is that Estimate the matrix for the intermediate unknown inertia Then the inertia matrix parameter adaptive law Specifically:

[0070]

[0071] Here, γ is a positive constant that can be designed.

[0072] The present invention also provides a computer device, comprising a memory and a processor, wherein the memory stores a computer program, and the processor implements the steps of the rigid spacecraft attitude tracking method as described above when executing the computer program.

[0073] The present invention also provides a computer-readable storage medium having a computer program stored thereon, which implements the steps of the rigid spacecraft attitude tracking method as described above when the computer program is executed by a processor.

[0074] The present invention also provides a computer program product, comprising a computer program, which implements the steps of the above-mentioned rigid spacecraft attitude tracking method when executed by a processor.

[0075] Compared with the prior art, the present invention has the following positive effects:

[0076] The rigid spacecraft attitude tracking method provided by the present invention introduces a preset time switching function as a time-varying gain to adjust the adaptive law of the attitude control torque and the uncertain inertia identification process, thereby ensuring that the spacecraft completes attitude tracking within a preset time and maintains tracking after the preset adjustment time, thereby improving the accuracy and reliability of control; when the inertia matrix parameters are unknown, the spacecraft can complete precise attitude tracking control within the preset time, ensure that the torque input remains continuous and bounded, avoid chattering, and maintain zero tracking error after the preset adjustment time, thereby having high engineering application value; before reaching the preset time, it can accurately complete precise tracking of the target attitude and accurately identify the true value of the spacecraft's inertia; at the preset time point, the control torque can avoid chattering; after the preset time point, the spacecraft system can maintain zero attitude tracking error, thereby improving the reliability and stability of the entire control system; this method does not require a complex parameter adjustment process, and the switching function structure is independent of the initial value of the system state, thereby greatly improving the compatibility and effectiveness of the control method and being adaptable to different spacecraft and working conditions. BRIEF DESCRIPTION OF THE DRAWINGS

[0077] The drawings in the specification are used to provide further understanding of the present invention and constitute a part of the present invention. The exemplary embodiments of the present invention and their descriptions are used to explain the present invention and do not constitute improper limitations on the present invention.

[0078] Figure 1 Schematic diagram of the process of the present invention;

[0079] Figure 2 This is a simulation result curve of the attitude quaternion tracking error of the first embodiment of the present invention;

[0080] Figure 3 1 is a simulation result curve of angular velocity tracking error according to an embodiment of the present invention;

[0081] Figure 4 The unknown inertia identification value of the first embodiment of the present invention is: and The simulation result curve of

[0082] Figure 5 The unknown inertia identification value of the first embodiment of the present invention is: and The simulation result curve of

[0083] Figure 6This is an input torque curve diagram of the first embodiment of the present invention. DETAILED DESCRIPTION

[0084] The present invention will be further described in detail below with reference to the accompanying drawings and specific embodiments, which are intended to explain the present invention rather than to limit it.

[0085] See also Figure 1 A rigid spacecraft attitude tracking method includes the following steps:

[0086] Step 1: Based on the attitude kinematics and dynamics equations of the rigid spacecraft body, the attitude kinematics and dynamics equations of the target rigid spacecraft are constructed to obtain the attitude quaternion error of the rigid spacecraft body relative to the target rigid spacecraft and the angular velocity error of the rigid spacecraft body coordinate system relative to the target rigid spacecraft coordinate system;

[0087] Step 2: Construct a switching function with respect to a preset time, establish a virtual controller based on the attitude quaternion error of the rigid spacecraft body relative to the target rigid spacecraft, and calculate the derivative of the virtual controller;

[0088] Step 3: Based on the attitude quaternion error of the rigid spacecraft body relative to the target rigid spacecraft and the angular velocity error of the rigid spacecraft body coordinate system relative to the target rigid spacecraft coordinate system, the three-axis control torque required to complete angular velocity tracking within a preset time is calculated in combination with the virtual controller and its derivatives;

[0089] Step 4: Perform matrix dimension deformation on the dynamic equation of the rigid spacecraft body to obtain the differential deformation equation; based on the three-axis control torque, perform time-invariant gain filtering on the variables in the differential deformation equation to obtain the first algebraic equation; perform time-invariant gain filtering on the variables in the first algebraic equation to obtain the second algebraic equation; perform time-varying gain filtering on the second algebraic equation to obtain the third algebraic equation;

[0090] Step 5: Based on the third algebraic equation, establish an adaptive law for the inertia matrix parameters of the rigid spacecraft body within a preset time, which is used to realize the inertia matrix parameter identification and realize the attitude tracking of the rigid spacecraft body.

[0091] The rigid spacecraft attitude tracking method provided by the present invention introduces a preset time switching function as a time-varying gain to adjust the adaptive law of the attitude control torque and the uncertain inertia identification process, thereby ensuring that the spacecraft completes attitude tracking within a preset time and maintains tracking after the preset adjustment time, thereby improving the accuracy and reliability of control; when the inertia matrix parameters are unknown, the spacecraft can complete precise attitude tracking control within the preset time, ensure that the torque input remains continuous and bounded, avoid chattering, and maintain zero tracking error after the preset adjustment time, thereby having high engineering application value; before reaching the preset time, it can accurately complete precise tracking of the target attitude and accurately identify the true value of the spacecraft's inertia; at the preset time point, the control torque can avoid chattering; after the preset time point, the spacecraft system can maintain zero attitude tracking error, thereby improving the reliability and stability of the entire control system; this method does not require a complex parameter adjustment process, and the switching function structure is independent of the initial value of the system state, thereby greatly improving the compatibility and effectiveness of the control method and being adaptable to different spacecraft and working conditions.

[0092] Preferably, step one is specifically:

[0093] The attitude kinematic equation of the rigid spacecraft body is:

[0094]

[0095] The attitude dynamics equation of the rigid spacecraft body is:

[0096]

[0097] in: is the attitude quaternion of the rigid spacecraft body coordinate system relative to the inertial reference coordinate system at time t, q0(t) is the scalar part of q(t), is the vector part of q(t), is the first-order derivative of q(t) with respect to time, is the angular velocity vector of the spacecraft body coordinate system relative to the inertial reference coordinate system at time t, ω1(t), ω2(t) and ω3(t) are all components of ω(t), is the first-order derivative of ω(t) with respect to time, is the unknown moment of inertia matrix of the rigid spacecraft body, which is a symmetric matrix, specifically:

[0098]

[0099] is the three-axis control torque applied to the rigid spacecraft body. For a given vector a × The antisymmetric matrix generated for vector a is:

[0100]

[0101] Construct the attitude kinematics and dynamics equations of the target rigid spacecraft:

[0102] The attitude kinematic equation of the target rigid spacecraft is:

[0103]

[0104] Among them, I3 is the 3×3 identity matrix, is the attitude quaternion of the target rigid spacecraft coordinate system relative to the rigid spacecraft body coordinate system at time t, where q d0 (t) is q d The scalar part of (t), q d The vector part of (t), q d (t) the first derivative with respect to time, is the angular velocity vector in the target rigid spacecraft coordinate system relative to the rigid spacecraft body coordinate system at time t, ω d1 (t),ω d2 (t) and ω d3 (t) are all ω d The component in (t), ω d (t) the first derivative with respect to time,

[0105] The attitude dynamics equation of the target rigid spacecraft is:

[0106]

[0107] definition is the attitude quaternion error of the rigid spacecraft body relative to the target rigid spacecraft at time t, where for The scalar part of for The vector part of Specifically:

[0108]

[0109] definition is the angular velocity error of the rigid spacecraft body coordinate system relative to the target rigid spacecraft coordinate system at time t, where and Both The weight in Specifically:

[0110]

[0111] in, is the transformation matrix from the target rigid spacecraft coordinate system to the rigid spacecraft body coordinate system at time t.

[0112] Preferably, the switching function K(t) is specifically:

[0113]

[0114] Among them, t f is the preset time; t is the time.

[0115] Preferably, the virtual controller α(t) is specifically:

[0116]

[0117] Among them, c1 is a normal constant that can be designed;

[0118] Derivatives of virtual controllers Specifically:

[0119]

[0120] in, is the first-order derivative of the switching function K(t) at time t, specifically:

[0121]

[0122] Preferably, the three-axis control torque for completing angular velocity tracking within a preset time is specifically:

[0123]

[0124] Among them, c2 is a normal number that can be designed. is the unknown inertia estimation matrix, Specifically:

[0125]

[0126] in, of They are the unknown inertia matrix J at time t. 11 , J 22 , J 33 , J 23 , J 13 , J 12 estimated value.

[0127] Preferably, step four specifically includes:

[0128] definition There exists any given vector This makes equation (15) valid. Equation (15) is specifically:

[0129] Ja=L(a)θ(15)

[0130] Among them, L(a) is the matrix about the given vector a, specifically:

[0131]

[0132] Combined with the attitude dynamics equation of the rigid spacecraft body, we get:

[0133]

[0134] make The differential deformation equation is obtained:

[0135]

[0136] The variables in the differential transformation equation are filtered with time-invariant gain, specifically:

[0137]

[0138] Among them, k1 is a normal constant that can be designed, ω f1 (t), W f1 (t) and u f1 (t) are the output signals of the filter after the time-invariant gain filtering of the variables in the differential transformation equation at time t, and and are u at time t f1 (t),ω f1 (t) and W f1 (t) the first derivative with respect to time;

[0139] Defining variables This gives the first algebraic equation:

[0140]

[0141] For variables in first algebraic equations and u f1 (t) Perform filtering with time-invariant gain, specifically:

[0142]

[0143] Where k2 is a normal number that can be designed. and u f2 (t) are all variables in the first algebraic equation at time t and u f1 (t) The output signal of the filter after time-invariant gain filtering, and and They are u f2 (t) The first derivative with respect to time, specifically:

[0144] We get equation (22):

[0145]

[0146] definition is time t The adjoint matrix of is time t The value of the determinant is:

[0147] The second algebraic equation (23) is obtained:

[0148] Y(t)=Δ(t)θ (23)

[0149] The second algebraic equation is filtered with a time-varying gain, where the time-varying gain is the square of Δ(t), i.e., Δ 2 (t), specifically:

[0150]

[0151] Among them, Δ f (t) is the output signal after the second algebraic equation is filtered by time-varying gain, and Δ f (t) is the adaptive gain of the spacecraft inertia matrix parameter identification algorithm, Y f (t) is the output signal of the filter after the second algebraic equation is filtered with time-varying gain, and and Δ is the time t f (t), Y f (t) is the first derivative with respect to time, and The third algebraic equation is obtained:

[0152] Y f (t) = Δ f (t)θ (25).

[0153] Preferably, Estimate the matrix for the intermediate unknown inertia Inertia matrix parameter adaptation law Specifically:

[0154]

[0155] Where γ is a positive constant that can be designed, The six variables in are the six elements J in the unknown inertia matrix J of the spacecraft. 11 ,J 22 ,J 33 ,J 23 ,J 13 ,J 12 The estimated values ​​are:

[0156]

[0157] In order to verify that the rigid spacecraft disclosed in the present invention can complete the accurate tracking of the preset time target under the condition of inertia uncertainty, the initial attitude quaternion value of the rigid spacecraft body is selected.

[0158] q v (0) = [0.4 -0.35 0.4] T , q0(0)=-0.7467, the initial angular velocity value of the rigid spacecraft body ω(0)=[0.2 0.2 0.2] T , unknown inertia matrix According to formula (1) and formula (2), the attitude kinematics and dynamics equations of the rigid spacecraft body are established; the initial attitude quaternion value q of the target rigid spacecraft is selected dv (0) = [0.2 -0.15 0.3] T ,q d0 (0) = 0.9206, the angular velocity of the target rigid spacecraft is set to

[0159] ω d (t)=[0.1sin(0.2πt) 0.1sin(0.2πt) 0.1sin(0.2πt)] T , the attitude kinematics and dynamics equations of the target rigid spacecraft are established according to formulas (5) and (6); the attitude quaternion error and angular velocity error of the rigid spacecraft body tracking the target rigid spacecraft are calculated according to formulas (7) and (8).

[0160] Preset time t f = 30 seconds, calculate the virtual controller α(t) according to formula (10), where the design parameter c1 = 50, and calculate the time derivative of the virtual controller α(t) according to formula (11)

[0161] The three-axis control torque u(t) of the rigid spacecraft is calculated according to formula (13), where the design parameter c2 = 100, and the variable W(t) in formula (18) is calculated.

[0162] According to formula (19), the angular velocity ω(t) and the three-axis control torque u(t) of the rigid spacecraft and the variable W(t) calculated in step 4 are filtered to obtain u f1 (t) and The gain of the time-invariant filter is k1 = 0.2.

[0163] According to formula (21), the filtering result u of step 5 is f1 (t) and Perform time-invariant filtering to obtain u f2 (t) and The gain of the time-invariant filter is k2 = 0.01, and the variables Y(t) and Δ(t) are calculated according to formulas (22) and (23).

[0164] According to formula (24), Y(t) and Δ(t) in step 6 are filtered to obtain Y f (t) and Δ f (t), where the gain of the time-varying filter is the square of Δ(t) in step 6, that is, Δ 2 (t).

[0165] According to formula (26), the adaptive law for the identification of the inertia parameters of rigid spacecraft is calculated. The parameter to be designed is γ=50000.

[0166] In the Matlab / Simulink environment, the embodiment 1 is simulated and verified. The simulation results are shown in Figures 2 to 6 ,in, Figure 2 is the spacecraft attitude quaternion tracking error curve, Figure 3 This is the spacecraft angular velocity tracking error curve. It can be seen that the spacecraft attitude system completes attitude tracking and angular velocity tracking within the preset adjustment time of 30 seconds, and maintains the zero error of the system tracking target after the preset time of 30 seconds; Figure 4 and Figure 5 This is the curve of the unknown inertia identification value of the spacecraft obtained by the inertia online identification method of the present invention. It can be seen that the identification method can identify the true value of the unknown inertia of the spacecraft in 30 seconds; Figure 6 The input torque of the spacecraft during the entire tracking process is continuous and bounded during the control process, and no vibration will occur at the preset time point of 30 seconds, which has high engineering application value.

[0167] Although this method does not know the spacecraft's moment of inertia matrix, it can further design the spacecraft's input torque through backstepping, introducing virtual control and the adaptive law of the moment of inertia parameters, enabling the spacecraft to complete target attitude tracking within a preset time. The preset time control method designed by the present invention differs from traditional preset time control methods, mainly in its flexibility and simplicity. The switching function in the preset time control is independent of the spacecraft's initial state value, eliminating the complex initial calculation process and the need to satisfy complex functional relationships. The controller structure is simple and does not need to rely on the spacecraft's true inertia value. Before reaching the preset time, the spacecraft can accurately complete the precise tracking of the target attitude and accurately identify the spacecraft's true inertia value; at the preset time point, the control torque can avoid chattering; after the preset time point, the spacecraft system can maintain zero attitude tracking error.

[0168] Based on the same inventive concept, an embodiment of the present application provides a computer device comprising a memory, a processor, and a computer program stored in the memory and executable on the processor. When the processor executes the computer program, the steps of a rigid spacecraft attitude tracking method are implemented. The memory may include internal memory, such as a high-speed random access memory, and may also include non-volatile memory, such as at least one disk storage device. The processor, network interface, and memory are interconnected via an internal bus. The internal bus may be an industrial standard architecture bus, a peripheral component interconnect standard bus, an extended industrial standard architecture bus, etc. The bus may be divided into an address bus, a data bus, a control bus, etc. The memory is used to store programs. Specifically, the program may include program code, and the program code includes computer operating instructions. The memory may include internal memory and non-volatile memory, and provides instructions and data to the processor.

[0169] Based on the same inventive concept, an embodiment of the present application provides a computer-readable storage medium, wherein the computer-readable storage medium stores a computer program. When the computer program is executed by a processor, the computer-readable storage medium implements the steps of the rigid spacecraft attitude tracking method. Specifically, the computer-readable storage medium includes, but is not limited to, volatile memory and / or non-volatile memory. The volatile memory may include RAM (Random Access Memory) and / or cache memory, etc. The non-volatile memory may include ROM (Read Only Memory), a hard disk, a flash memory, an optical disk, a magnetic disk, etc.

[0170] Based on the same inventive concept, an embodiment of the present application provides a computer program product, which includes a computer program stored on a computer-readable storage medium, and the computer program includes program instructions. When the program instructions are executed by a computer device, the computer device executes the steps of the above-mentioned rigid spacecraft attitude tracking method.

[0171] Those skilled in the art will appreciate that embodiments of the present invention may be provided as methods, systems, or computer program products. Thus, the present invention may take the form of an entirely hardware embodiment, an entirely software embodiment, or an embodiment combining software and hardware aspects. Furthermore, the present invention may take the form of a computer program product implemented on one or more computer-usable storage media (including but not limited to magnetic disk storage, CD-ROM (Compact Disc Read-Only Memory), optical storage, etc.) containing computer-usable program code.

[0172] Finally, it should be noted that the embodiments listed above are merely one or more specific manifestations of the technical solution of the present invention. Their purpose is to clearly illustrate the concept, principles, and application of the present invention through specific examples, and is in no way intended to limit the scope of protection of the present invention to these specific embodiments. In fact, the true value of this invention lies in its technical ideas and innovations, not in its form of expression or implementation.

Claims

1. A method for tracking the attitude of a rigid spacecraft, characterized in that: The following steps are involved: Step 1: Based on the attitude kinematics and dynamics equations of the rigid spacecraft body, the attitude kinematics and dynamics equations of the target rigid spacecraft are constructed to obtain the attitude quaternion error of the rigid spacecraft body relative to the target rigid spacecraft and the angular velocity error of the rigid spacecraft body coordinate system relative to the target rigid spacecraft coordinate system; Step 2: Construct a switching function with respect to a preset time, establish a virtual controller based on the attitude quaternion error of the rigid spacecraft body relative to the target rigid spacecraft, and calculate the derivative of the virtual controller; Step 3: Based on the attitude quaternion error of the rigid spacecraft body relative to the target rigid spacecraft and the angular velocity error of the rigid spacecraft body coordinate system relative to the target rigid spacecraft coordinate system, the three-axis control torque required to complete angular velocity tracking within a preset time is calculated in combination with the virtual controller and its derivatives; Step 4: Perform matrix dimension deformation on the dynamic equation of the rigid spacecraft body to obtain the differential deformation equation; based on the three-axis control torque, perform time-invariant gain filtering on the variables in the differential deformation equation to obtain the first algebraic equation; perform time-invariant gain filtering on the variables in the first algebraic equation to obtain the second algebraic equation; perform time-varying gain filtering on the second algebraic equation to obtain the third algebraic equation; Step 5: Based on the third algebraic equation, establish an adaptive law for the inertia matrix parameters of the rigid spacecraft body within a preset time, which is used to realize the inertia matrix parameter identification and realize the attitude tracking of the rigid spacecraft body.

2. A rigid spacecraft attitude tracking method according to claim 1, characterized in that: In step 1, the attitude kinematic equation of the rigid spacecraft body is: Among them, q(t) is the attitude quaternion of the rigid spacecraft body coordinate system relative to the inertial reference coordinate system at time t, q0(t) is the scalar part of q(t), and q v (t) is the vector part of q(t), is the first-order derivative of q(t) with respect to time, ω(t) is the angular velocity vector of the spacecraft body coordinate system relative to the inertial reference coordinate system at time t, and I3 is the unit matrix; The attitude dynamics equation of the rigid spacecraft body is: in, is the first-order derivative of ω(t) with respect to time, J is the unknown moment of inertia matrix of the rigid spacecraft body, which is a symmetric matrix; u(t) is the three-axis control torque applied to the rigid spacecraft body; The attitude kinematic equation of the target rigid spacecraft is: Among them, q d (t) is the attitude quaternion of the target rigid spacecraft coordinate system relative to the rigid spacecraft body coordinate system at time t, where q d0 (t) is q d The scalar part of (t), q dv (t) is q d The vector part of (t), q d (t) The first derivative with respect to time, ω d (t) is the angular velocity vector in the target rigid spacecraft coordinate system relative to the rigid spacecraft body coordinate system at time t; The attitude dynamics equation of the target rigid spacecraft is: in, ω d (t) the first derivative with respect to time; Attitude quaternion error of the rigid spacecraft body relative to the target rigid spacecraft for: Angular velocity error of the rigid spacecraft body coordinate system relative to the target rigid spacecraft coordinate system for: in, is the transformation matrix from the target rigid spacecraft coordinate system to the rigid spacecraft body coordinate system at time t.

3. A rigid spacecraft attitude tracking method according to claim 2, characterized in that: The switching function K(t) is: Among them, t f is the preset time.

4. A rigid spacecraft attitude tracking method according to claim 3, characterized in that: The virtual controller α(t) is: Among them, c1 is a normal constant that can be designed; for The vector part of Derivatives of virtual controllers for: in, for The scalar part of is the first-order derivative of the switching function K(t) at time t, specifically:

5. A rigid spacecraft attitude tracking method according to claim 4, characterized in that: The three-axis control torque u(t) required to complete angular velocity tracking within the preset time is calculated as follows: Among them, c2 is a normal number that can be designed. is the unknown inertia estimation matrix, Specifically: in, of They are the unknown inertia matrix J at time t. 11 , J 22 , J 33 , J 23 , J 13 , J 12 estimated value.

6. A method for tracking the attitude of a rigid spacecraft according to claim 5, characterized in that: In step 4, let [J 11 J 22 J 33 J 23 J 13 J 12 ] T is the intermediate matrix θ, and the differential transformation equation is: Among them, W(t) is the intermediate parameter, specifically: W(t)=-ω × (t)L(ω(t)) Where, L(ω(t)) is the matrix about ω(t); The variables in the differential transformation equation are filtered with time-invariant gain, specifically: Among them, k1 is a normal constant that can be designed, ω f1 (t), W f1 (t) and u f1 (t) are the output signals of the filter after the time-invariant gain filtering of the variables in the differential transformation equation at time t, and are u at time t f1 (t),ω f1 (t) and W f1 (t) the first derivative with respect to time; Set variable The first algebraic equation is: For variables in first algebraic equations and u f1 (t) Perform filtering with time-invariant gain, specifically: Where k2 is a normal number that can be designed. and u f2 (t) are all variables in the first algebraic equation at time t and u f1 (t) performing time-invariant gain filtering on the output signal of the filter; and They are u f2 (t) The first derivative with respect to time, specifically: The equation is: definition is time t The adjoint matrix of is time t The value of the determinant is: The second algebraic equation is obtained: Y(t)=Δ(t)θ The second algebraic equation is filtered with a time-varying gain, where the time-varying gain is the square of Δ(t), i.e., Δ 2 (t), specifically: Among them, Δ f (t) is the output signal after the second algebraic equation is filtered by time-varying gain, and Δ f (t) is the adaptive gain of the spacecraft inertia matrix parameter identification algorithm, Y f (t) is the output signal of the filter after time-varying gain filtering of the second algebraic equation; and Δ is the time t f (t), Y f The first-order derivative of (t) with respect to time gives the third algebraic equation: Y f (t)=Δ f (t)θ.

7. A method for tracking the attitude of a rigid spacecraft according to claim 6, characterized in that: set up Estimate the matrix for the intermediate unknown inertia Then the inertia matrix parameter adaptive law Specifically: Here, γ is a positive constant that can be designed.

8. A computer device comprising a memory and a processor, wherein the memory stores a computer program, wherein: When the processor executes the computer program, the steps of the rigid spacecraft attitude tracking method according to any one of claims 1 to 7 are implemented.

9. A computer-readable storage medium having a computer program stored thereon, characterized in that: When the computer program is executed by a processor, the steps of the rigid spacecraft attitude tracking method according to any one of claims 1 to 7 are implemented.

10. A computer program product comprising a computer program, characterized in that When the computer program is executed by a processor, the steps of the rigid spacecraft attitude tracking method according to any one of claims 1 to 7 are implemented.

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