Satellite attitude maneuvering method after specific failure reconstruction of a five-pyramid configuration sgcmg

Abstract

A satellite attitude maneuvering method after specific failure reconstruction of a pentagonal pyramidal SGCMG configuration is disclosed, belonging to the field of spacecraft attitude control technology. This includes a single-axis pointing maneuvering strategy for satellites equipped with a pentagonal pyramidal configuration and requiring single-axis spatial pointing. The designed method ensures that the angular momentum of the control moment gyroscope combination is in an optimal state under the target observation attitude by designing the initial angular momentum configuration and target attitude. The target maneuver is divided into two stages: singularity avoidance and target maneuver, ensuring that the angular momentum of the control moment gyroscope combination does not become singular during the maneuver. The maneuvering strategy proposed in this invention guarantees that the satellite does not experience angular momentum singularity during the target maneuver and that the angular momentum is in an optimal state during target observation. For satellites that lose three-axis maneuvering capability due to partial failure of the control moment gyroscope, this method ensures their mission execution capability and improves satellite availability.

Description

Technical Field

[0001] This invention relates to a satellite attitude maneuvering method after specific failure reconstruction of a pentagonal pyramidal SGCMG, belonging to the field of spacecraft attitude control technology. Background Technology

[0002] With the emergence of increasingly complex and diverse observation missions, optical satellites are required to be able to point at any space target in real time. That is, the satellite must control its attitude to point a certain vector (e.g., satellite + Z-axis) at any target in space. The control moment gyroscope (SGCMG) can output significant control torque during attitude maneuvering, has a fast dynamic response, and can well meet the attitude maneuvering torque requirements of agile satellites, thus gaining widespread application.

[0003] From a satellite safety control perspective, if several control moment gyroscopes in a control moment gyroscope assembly fail, the remaining intact control moment gyroscopes are reconfigured to form a new control moment gyroscope assembly. If the reconfigured control moment gyroscope group contains at least four gyroscopes, the satellite generally retains some three-axis maneuverability. However, if the effective number of gyroscopes is less than four, the control moment gyroscopes will have small angular momentum envelopes and contain numerous complex singularities, resulting in a significant reduction in maneuverability and the loss of three-axis maneuverability. Summary of the Invention

[0004] The technical problem solved by this invention is to overcome the shortcomings of existing technologies and provide a satellite attitude maneuvering method after specific failure reconstruction of a pentagonal pyramidal SGCMG configuration. By configuring the initial angular momentum and designing a suitable target observation attitude, the satellite's angular momentum is kept at the center of the ring during target observation, ensuring the satellite is in its safest state. A distributed maneuvering strategy is designed to allow the satellite to maneuver to the target attitude in two steps, avoiding singularities in the combined angular momentum of the control moment gyroscope.

[0005] The technical solution of this invention is: a satellite attitude maneuvering method after specific failure reconstruction of a pentagonal pyramidal SGCMG, comprising:

[0006] After receiving the target pointing mission, calculate the initial attitude of the inertial guarding system, and rotate the guarding attitude and its position around the satellite body coordinate system + Z axis by any angle α. rotZ The attitude is used as the target observation attitude;

[0007] Calculate the optimal initial bias angular momentum based on the inertial conservation initial attitude;

[0008] Calculate the target observation attitude based on the optimal initial bias angular momentum;

[0009] The target maneuver is decomposed into two steps: around the singular axis and parallel to the singular axis. The singular avoidance attitude after maneuvering around the singular axis is calculated. The singular avoidance attitude is used as the target attitude for the first maneuver. After the maneuver is completed, the target observation attitude is used as the target attitude for the second maneuver.

[0010] Furthermore, the initial attitude of the inertial guarding is

[0011] Where sAngle is the maneuver angle, sAxis is the maneuver axis, and q ib This indicates the current satellite attitude.

[0012] Furthermore, the method for calculating the optimal initial bias angular momentum based on the inertial conservation initial attitude is as follows: the initial bias angular momentum calculation is converted into a nonlinear programming problem, which is solved using a particle swarm optimization algorithm; the nonlinear programming model is...

[0013]

[0014] in, Let be the initial angular momentum. Let n be the angular momentum of the control moment gyroscope group at the current attitude. qy denoted by angular momentum singularity axis, and sd is the optimal toroidal designation.

[0015] Furthermore, the target observation attitude is Where, q it_int This is the initial attitude for inertial guarding.

[0016] Furthermore, the bias angular momentum is calculated using an optimal control algorithm. Ensure that the satellite's angular momentum remains within a safe region of the annular cavity under both the initial attitude and the target observation attitude.

[0017] Furthermore, the singular avoidance posture is Where, △q jd To avoid maneuvering quaternions, the satellite rotates around a singular axis to ensure that the satellite's angular momentum does not become singular.

[0018] Furthermore, the target attitude after the second maneuver is a guarded attitude, ensuring that the satellite does not experience any singularities during the second maneuver and that the target is within the satellite's observation range after the maneuver is completed.

[0019] Furthermore, the target's angular velocity does not exceed the angular velocity limit value throughout the entire maneuver.

[0020] A computer-readable storage medium storing a computer program, which, when executed by a processor, implements the steps of a satellite attitude maneuvering method after specific failure reconstruction of a pentagonal pyramidal SGCMG configuration.

[0021] A satellite attitude maneuvering device after specific failure reconstruction of a pentagonal pyramidal SGCMG configuration includes a memory, a processor, and a computer program stored in the memory and executable on the processor. When the processor executes the computer program, it implements the steps of the satellite attitude maneuvering method after specific failure reconstruction of a pentagonal pyramidal SGCMG configuration.

[0022] The advantages of this invention compared to the prior art are:

[0023] (1) This invention configures the initial bias angular momentum through the optimal control algorithm, so that the angular momentum of the target observation attitude is on the optimal circle, thus maximizing the observation capability.

[0024] (2) By dividing the target maneuver into two segments, the present invention allows the satellite to maneuver to the target attitude in two steps, thus ensuring that the angular momentum does not become singular during the target maneuver. Attached Figure Description

[0025] Various other advantages and benefits will become apparent to those skilled in the art upon reading the following detailed description of preferred embodiments. The accompanying drawings are for illustrative purposes only and are not intended to limit the invention. Furthermore, the same reference numerals denote the same parts throughout the drawings. In the drawings:

[0026] Figure 1 This is a schematic diagram of the installation of the pentagonal pyramidal control torque gyroscope of the present invention;

[0027] Figure 2 This is a schematic diagram of the singular surface after the failure reconstruction of the pentagonal pyramid control torque gyroscope combination of the present invention;

[0028] Figure 3 This is a schematic diagram of the method flow of the present invention. Detailed Implementation

[0029] To better understand the above technical solutions, the technical solutions of the present invention will be described in detail below with reference to the accompanying drawings and specific embodiments. It should be understood that the embodiments of the present invention and the specific features in the embodiments are detailed descriptions of the technical solutions of the present invention, rather than limitations on the technical solutions of the present invention. In the absence of conflict, the embodiments of the present invention and the technical features in the embodiments can be combined with each other.

[0030] The following description, in conjunction with the accompanying drawings, provides a more detailed explanation of the satellite attitude maneuvering method after specific failure reconstruction of a pentagonal pyramidal SGCMG configuration provided by the embodiments of the present invention. Figure 3 Specific implementation methods may include:

[0031] Given that the current +Z axis points to OZ0 and the satellite's current attitude qib The satellite received a target pointing instruction, pointing to target OZ1:

[0032] Step 1: After receiving the new task, calculate the guarding posture q. it_int Then the guard attitude and its attitude when rotated around the +Z axis at any angle can be used as the target observation attitude;

[0033] Step 2: Based on the defensive posture, design a nonlinear programming model to calculate the optimal initial bias angular momentum;

[0034] Step 3: Calculate the target observation attitude q based on the optimal initial bias angular momentum. it ;

[0035] Step 4: Decompose the target maneuver into two steps: maneuvering around the singular axis and maneuvering parallel to the singular axis. Calculate the singular avoidance attitude q after maneuvering around the singular axis. qy The singular evasion attitude is used as the target attitude for the first maneuver. After the maneuver is completed, the target observation attitude is used as the target attitude for the next maneuver. The angular velocity does not exceed the angular velocity limit value throughout the entire maneuver.

[0036] In the solutions provided in the embodiments of the present invention, such as Figure 1 When the pentagonal pyramidal control moment gyroscope assembly has only 3 units remaining on its sides ( Figure 1 In the control moment gyroscope (green indicates effective, red indicates ineffective), when effective, the angular momentum singular breadth of the three CMGs is as follows: Figure 2 As shown, singularities accumulate within the region near angular momentum 0, providing virtually no angular momentum for satellite maneuverability. At the center of angular momentum lies a tilted, columnar singular region that tapers from the middle to the ends. Surrounding this columnar region is a ring-shaped cavity at approximately 37° to the XY plane. The ring has an inner diameter of 0.14h0 Nms, an outer diameter of 1.1h0 Nms, and a radius of 0.48h0 Nms, where h0 is the angular momentum of a single control moment gyroscope. The cavity is surrounded by six helical singular surfaces. In steady state, the angular momentum of the control moment gyroscope assembly is biased towards the central ring of the ring. Figure 2 (Middle red line) indicates the safest state for the satellite. The angular momentum required for satellite maneuvering must be within a toroidal cavity, i.e., rotating around a singular cylinder; otherwise, the control moment gyroscope assembly cannot provide the torque required for maneuvering. For this specific reconfiguration state, this invention designs an attitude maneuvering method. By configuring the initial angular momentum and designing a suitable target observation attitude, the satellite's angular momentum is kept at the center of the toroidal cavity during target observation, ensuring the satellite is in its safest state. A distributed maneuvering strategy is designed, allowing the satellite to maneuver to the target attitude in two steps, avoiding singularity in the angular momentum of the control moment gyroscope assembly.

[0037] 1) Calculate feasible target observation attitudes

[0038] The target space in the inertial frame that the satellite is pointing to is OZ1; the pointing vector of the current satellite's +Z axis in the inertial frame is OZ0; and the current satellite attitude is q. ib Using the shortest path maneuver, rotating around a maneuver axis that is perpendicular to both OZ1 and OZ0, the initial attitude q of inertial protection is obtained. it_int0 The maneuver axis and maneuver angle are as follows:

[0039]

[0040] sAngle = arccos(OZ0) T ·OZ1)

[0041] in For quaternion multiplication, M(※) is the attitude rotation matrix corresponding to ※.

[0042] In the initial inertial control attitude, the satellite's +Z axis is already pointing towards the target. Based on this, the satellite rotates its +Z axis by any angle α. rotZ The posture after the target can be used for target pointing and can be used as the posture for target observation.

[0043] 2) Calculate the initial bias angular momentum

[0044] If the angular momentum of the control moment gyroscope group under the current attitude is (By default, the angular momentum is on the optimal ring in the current attitude). Magnetic unloading can configure the angular momentum to a suitable position, i.e., the initial angular momentum. This ensures that, under the target observation attitude, the angular momentum is located on the center line of the annular cavity.

[0045] Constraint 1: Initial angular momentum It is located inside an annular cavity.

[0046]

[0047] Where n qy The direction of the singular axis of angular momentum.

[0048] Constraint 2: By setting angle α rotZ Then, the control torque gyro angular momentum under the target observation attitude is also on the optimal ring.

[0049] Calculate the control torque gyroscope bias angular momentum under the initial reference of inertial conservation:

[0050]

[0051]

[0052] To ensure that the angular momentum of the control torque gyroscope under the target's observed attitude is also on the optimal ring, the optimal ring marker sd must meet the following requirements:

[0053]

[0054] In addition to satisfying the above two constraints, it is also required that the angular momentum requiring magnetic unloading be minimized, that is, the deviation between the initial angular momentum and the current angular momentum be minimized:

[0055]

[0056] The initial bias angular momentum calculation is then transformed into a nonlinear programming problem, which is solved using the particle swarm optimization algorithm. The nonlinear programming model is as follows:

[0057]

[0058] 3) Calculate the target observation attitude

[0059] θ tp1 =arcsin(sd)

[0060] θ tp2 =p-θ tp1

[0061] α tp1 =θ tp1 -atan2(sa,sb)

[0062] α tp2 =θ tp2 -atan2(sa,sb)

[0063] α tp1 =mod(α) tp1 +p,2π)-p

[0064] α tp2 =mod(α) tp2 +p,2π)-p

[0065] When |sd|=0, the two angles are equal; when they are not equal, the smaller rotation angle is taken as the final rotation angle around the Z-axis.

[0066]

[0067] The target observation attitude quaternion is:

[0068]

[0069] 4) Design segmented maneuver strategies

[0070] The above design ensures that the satellite is in an optimal angular momentum state during target observation. However, during maneuvers, the satellite's angular momentum may still cross the central columnar singularity region. Further design of the maneuver process is needed.

[0071] During the transition from the current attitude to the inertial guarded attitude, first perform columnar singularity avoidance, and then... Angular momentum at target attitude after bypassing the columnar singularity region On the same side, move to The rotation principle is: around n qy Rotation angle θ jd Make the angular momentum vector Go to n qy and On the plane in which it is located. Calculation of the target quaternion for singularity surface avoidance:

[0072]

[0073] Angle of avoidance around the singular axis:

[0074]

[0075] Calculate the rotation axes of the angular momentum vectors of the two reference frames:

[0076]

[0077] n1 = sign(n qyc ·n H )·n qyc

[0078] Calculate the quaternion Δq for evasive maneuvers jd :

[0079]

[0080] The target maneuver is divided into two steps. The first step is to use q it_jd The target quaternion is q. After the maneuver reaches its target position, the second maneuver is performed, using the target's inertial defensive attitude as the target attitude for the second maneuver. it_int .

[0081] During the maneuver, the angular velocity is limited, and the maximum angular velocity does not exceed ω. max =0.48h0 / J sat J sat This is the moment of inertia of the satellite's principal axis of maximum inertia.

[0082] This invention provides a computer-readable storage medium storing computer instructions that, when executed on a computer, cause the computer to perform... Figure 3The method described.

[0083] Those skilled in the art will understand that embodiments of the present invention can be provided as methods, systems, or computer program products. Therefore, the present invention can take the form of a completely hardware embodiment, a completely software embodiment, or an embodiment combining software and hardware aspects. Furthermore, the present invention can take the form of a computer program product embodied on one or more computer-usable storage media (including, but not limited to, disk storage and optical storage) containing computer-usable program code.

[0084] This invention is described with reference to flowchart illustrations and / or block diagrams of methods, apparatus (systems), and computer program products according to embodiments of the invention. It will be understood that each block of the flowchart illustrations and / or block diagrams, and combinations of blocks in the flowchart illustrations and / or block diagrams, can be implemented by computer program instructions. These computer program instructions can be provided to a processor of a general-purpose computer, special-purpose computer, embedded processor, or other programmable data processing apparatus to produce a machine, such that the instructions, which execute via the processor of the computer or other programmable data processing apparatus, generate instructions for implementing the flowchart illustrations and / or block diagrams. Figure 1 One or more processes and / or boxes Figure 1 A device that provides the functions specified in one or more boxes.

[0085] These computer program instructions may also be stored in a computer-readable storage medium that can direct a computer or other programmable data processing device to function in a particular manner, such that the instructions stored in the computer-readable storage medium produce an article of manufacture including instruction means, which are implemented in a process Figure 1 One or more processes and / or boxes Figure 1 The function specified in one or more boxes.

[0086] These computer program instructions may also be loaded onto a computer or other programmable data processing equipment to cause a series of operational steps to be performed on the computer or other programmable equipment to produce a computer-implemented process, thereby providing instructions that execute on the computer or other programmable equipment for implementing the process. Figure 1 One or more processes and / or boxes Figure 1 The steps of the function specified in one or more boxes.

[0087] Obviously, those skilled in the art can make various modifications and variations to this invention without departing from its spirit and scope. Therefore, if these modifications and variations fall within the scope of the claims of this invention and their equivalents, this invention also intends to include these modifications and variations.

[0088] The contents not described in detail in this specification are common knowledge to those skilled in the art.

Claims

1. A satellite attitude maneuvering method after specific failure reconstruction in a pentagonal pyramidal SGCMG configuration, characterized in that, include: After receiving the target pointing mission, calculate the initial attitude of the inertial guarding system, and rotate the guarding attitude and its position around the satellite body coordinate system + Z-axis by any angle. The attitude is used as the target observation attitude; Calculate the optimal initial bias angular momentum based on the inertial conservation initial attitude; Calculate the target observation attitude based on the optimal initial bias angular momentum; The target maneuver is decomposed into two steps: around the singular axis and parallel to the singular axis. The singular avoidance attitude after maneuvering around the singular axis is calculated. The singular avoidance attitude is used as the target attitude for the first maneuver. After the maneuver is completed, the target observation attitude is used as the target attitude for the second maneuver. The method for calculating the optimal initial bias angular momentum based on the inertial guarded initial attitude is as follows: the initial bias angular momentum calculation is converted into a nonlinear programming problem, which is solved using a particle swarm optimization algorithm; the nonlinear programming model is... in, Let be the initial angular momentum. The angular momentum of the control moment gyroscope group at the current attitude. The direction of the singular axis of angular momentum The best circular symbol.

2. The satellite attitude maneuvering method after specific failure reconstruction of a pentagonal pyramidal SGCMG according to claim 1, characterized in that, The initial attitude of the inertial guarding is ;in, For the maneuver angle, For motor shaft, This indicates the current satellite attitude.

3. The satellite attitude maneuvering method after specific failure reconstruction of a pentagonal pyramidal SGCMG according to claim 1, characterized in that, The target observation attitude is: ;in, This is the initial attitude for inertial guarding.

4. The satellite attitude maneuvering method after specific failure reconstruction of a pentagonal pyramidal SGCMG according to claim 1, characterized in that, The bias angular momentum is calculated using the optimal control algorithm. This ensures that the satellite's angular momentum remains within a safe region of the annular cavity under both the initial attitude and the target observation attitude.

5. The satellite attitude maneuvering method after specific failure reconstruction of a pentagonal pyramidal SGCMG according to claim 1, characterized in that, The strange evasion posture is ,in, To avoid maneuvering quaternions, the satellite rotates around a singular axis to ensure that the satellite's angular momentum does not become singular.

6. The satellite attitude maneuvering method after specific failure reconstruction of a pentagonal pyramidal SGCMG according to claim 1, characterized in that, The target attitude after the second maneuver is a guarded attitude, ensuring that the satellite does not experience any singularities during the second maneuver and that the target is within the satellite's observation range after the maneuver is completed.

7. The satellite attitude maneuvering method after specific failure reconstruction of a pentagonal pyramidal SGCMG according to claim 1, characterized in that, The target's angular velocity does not exceed the angular velocity limit throughout the entire maneuver.

8. A computer-readable storage medium storing a computer program, characterized in that, When the computer program is executed by a processor, it implements the steps of the method as described in any one of claims 1 to 7.

9. A satellite attitude maneuvering device after specific failure reconstruction of a pentagonal pyramidal SGCMG, comprising a memory, a processor, and a computer program stored in the memory and executable on the processor, characterized in that: When the processor executes the computer program, it implements the steps of the method as described in any one of claims 1 to 7.

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