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Flexible satellite attitude orbit coupling control method based on isolation allowance method and pulse width fusion strategy

A flexible satellite, pulse width fusion technology, applied in the direction of the guidance device of the aerospace vehicle, etc., can solve the problem of not giving the thruster layout, not considering the pull of the plume's rotational inertia, and the flywheel being uncontrollable.

Active Publication Date: 2015-05-06
HARBIN INST OF TECH
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  • Summary
  • Abstract
  • Description
  • Claims
  • Application Information

AI Technical Summary

Problems solved by technology

[0007] The purpose of the present invention is to solve the problem that two sets of actuators for orbit control and attitude control need to be equipped during the attitude and orbit control process of the satellite, and the flywheel will not be able to control the layout of the corresponding thrusters without consideration of the plume. Due to the problems caused by the pulling deviation of the influence and moment of inertia, the failure to consider the isolation margin and the failure of the attitude to meet the requirements and cause serious losses, a flexible satellite attitude-orbit coupling control method based on the isolation margin method and the pulse width fusion strategy is proposed.

Method used

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  • Flexible satellite attitude orbit coupling control method based on isolation allowance method and pulse width fusion strategy
  • Flexible satellite attitude orbit coupling control method based on isolation allowance method and pulse width fusion strategy
  • Flexible satellite attitude orbit coupling control method based on isolation allowance method and pulse width fusion strategy

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specific Embodiment approach 1

[0068] Specific implementation mode 1: A flexible satellite attitude-orbit coupling control method based on the isolation margin method and the pulse width fusion strategy in this embodiment is specifically prepared according to the following steps:

[0069] Step 1. According to the geocentric inertial coordinate system (Oi, Xi, Yi, Zi) (ECI) such as Figure 21 , the satellite body coordinate system (Ob,Xb,Yb,Zb) such as Figure 14 , Satellite layout coordinate system (O1, X1, Y1, Z1) (the origin is taken at the geometric center of the satellite-array separation plane, the O1Z1 axis is within the satellite-array separation plane, pointing vertically to the star to the ground; the O1, Y1 axis is perpendicular to the satellite-array separation plane , pointing to the payload compartment; the O1X1 axis and the other two axes form the right-hand rule), and consider the influence of various types of disturbance torques on the attitude of various flexible satellites to conduct mathe...

specific Embodiment approach 2

[0080] Specific embodiment 2: The difference between this embodiment and specific embodiment 1 is that the disturbance torque of the rotating part described in step 1 is as follows Figure 16 and Figure 17 for:

[0081]

[0082] ω wby Indicates the rotational speed of the rotating part, I wby is the component of the inertia matrix of the rotating part relative to the origin; the disturbance torque T of the rotating part wb T wbx , T wby , T wbz Respectively in the corresponding x, y, z axis components;

[0083] The internal disturbance torque formula of the disturbance torque caused by the unlocking of the pyrotechnic device to the body is as follows Figure 18 ;

[0084]

[0085] ω wbz Indicates the rotational speed of the pyrotechnic product, I wbz is the component of the inertia matrix of the pyrotechnic product relative to the origin;

[0086] The separation torque of the separation body includes the small satellite separation disturbance torque such as ...

specific Embodiment approach 3

[0087] Embodiment 3: The difference between this embodiment and Embodiment 1 or 2 is that in step 1, the dynamic model of sailboard locking and satellite not being controlled is established. The specific process is as follows:

[0088] (1) Establish the attitude dynamic equation of the satellite with flexible solar panel attachment as:

[0089]

[0090] in,

[0091] Modal coordinates of sailboard A;

[0092] The rotational angular velocity of sailboard A;

[0093] R sa : Inertial dyadic for coupling sailboard rotation and whole star rotation;

[0094] f s : is the rotation coupling coefficient matrix of the sailboard A vibration to the whole star relative to the satellite itself;

[0095] R as : Inertia dyadic for coupling the rotation of the whole star and the rotation of the sailboard;

[0096] f A : is the rotational coupling coefficient matrix of the sailboard A vibration to the sailboard relative to the satellite system;

[0097] I a : Windsurfing A rel...

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Abstract

The invention relates to the field of flexible satellite attitude orbit coupling control, in particular to a flexible satellite attitude orbit coupling control method based on an isolation allowance method and a pulse width fusion strategy. The method solves the problems that in the satellite in-orbit attitude and orbit control process, a flywheel cannot control that the corresponding thruster layout is not given, the plume influence and rotational inertia level bias are not considered, isolation allowance is not considered and the attitude does not meet the requirement. The method comprises the steps of 1, obtaining sailboard locking and satellite uncontrollable dynamical model parameters; 2, determining installation position coordinates of a thrust; 3, determining an IM value; 4, obtaining an orbit-control LQG sequence; 5, determining the orbit-control pulse width and the air injection direction; 6, selecting air injection of the attitude-controlled thrust; 7, determining the range of the attitude-controlled thrust; 8, determining the attitude-controlled air injection time; 9, obtaining an equivalent force moment value. The flexible satellite attitude orbit coupling control method is applied to the field of flexible satellite attitude orbit coupling control.

Description

technical field [0001] The invention relates to an orbit coupling control method, in particular to an orbit coupling control method with a flexible satellite attitude. Background technique [0002] For the attitude and orbit control of satellites in orbit, people began to adopt the method of separate control of attitude and orbit "Integrated Coupling Control of Formation Satellite Relative Orbit and Attitude". If there is no set of actuators, the number of actuators such as thrusters will be increased, resulting in a waste of resources. In terms of modeling, the separation of attitude and orbit simplifies the modeling problem, but increases the complexity of the model itself, making the spacecraft control algorithm complex and cumbersome, and occupying the limited computing resources of the on-board computer. With the development of aerospace technology, in the face of new space missions that have attitude-orbit coupling problems such as spacecraft rendezvous and docking, a...

Claims

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Application Information

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Patent Type & Authority Applications(China)
IPC IPC(8): B64G1/24
Inventor 孙延超刘萌萌马广富王晓东李传江朱津津
Owner HARBIN INST OF TECH
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