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Decoupling control method for relative orbits and attitudes of formation satellites

A technology of relative orbit and decoupling control, applied in attitude control, three-dimensional position/course control, etc., can solve the problems of high control dimension, large calculation amount and low solution efficiency of formation satellites, so as to improve solution efficiency and reduce control The effect of dimensionality

Inactive Publication Date: 2010-08-04
HARBIN INST OF TECH
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Problems solved by technology

[0003] In order to solve the problem that the relative orbit and attitude of formation satellites are seriously coupled, resulting in high control dimension of formation satellites, resulting in a large amount of computation on the satellite and low solution efficiency, the present invention provides a decoupling control method for relative orbit and attitude of formation satellites

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  • Decoupling control method for relative orbits and attitudes of formation satellites
  • Decoupling control method for relative orbits and attitudes of formation satellites
  • Decoupling control method for relative orbits and attitudes of formation satellites

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specific Embodiment approach 1

[0013] Specific implementation mode one: according to the instructions attached figure 1 and 3 Specifically explaining this embodiment, the decoupling control process of the relative orbit and attitude decoupling control method of a formation satellite described in this embodiment is:

[0014] Step 1, analyzing the initial condition constraints and terminal condition constraints of the formation satellites composed of N satellites, the initial condition constraints include initial relative position, velocity, attitude and attitude angular velocity, and the terminal condition constraints include terminal relative position, velocity and attitude angular velocity;

[0015] Step 2, make the formation satellite meet the following conditions:

[0016] Condition 21. At adjacent moments, the change angle of the thrust vector direction of each slave satellite in the formation satellite is consistent with the attitude control capability of the satellite;

[0017] Condition 22. At any...

specific Embodiment approach 2

[0021] Specific embodiment two: This embodiment is a further description of the relative orbit and attitude decoupling control method of the formation satellites described in the specific embodiment one. The decoupling constraints of thrust vector maneuverability are: a T (k)a(k-1) / (‖a(k)‖·‖a(k-1)‖)≥cos(γ·ω max Δt), that is, the angle between the thrust vectors of each satellite in the formation satellite at two adjacent moments is less than or equal to the maximum angle achieved by the attitude maneuver of each satellite in the formation satellite, where a(k-1) and a(k) respectively represents the thrust acceleration vector of the slave star at k-1 time and k time, k is a natural number, ω max is the maximum attitude angular velocity achieved by the satellite, Δt is the control sampling time, and the thrust vector maneuver constraint weight coefficient γ∈(0,1).

[0022] The thrust vector maneuver capability decoupling constraint described in this embodiment is the thrust ve...

specific Embodiment approach 3

[0023] Embodiment 3: This embodiment is a further description of the relative orbit and attitude decoupling control method of the formation satellites described in Embodiment 1. The angle decoupling constraint between the star sensor optical axis 1 vector and the thrust vector is: from the star sensor optical axis 1 vector v cam with thrust vector v th The angle β between them is greater than or equal to v sun with v cam The minimum allowable angle α between, and is less than or equal to the supplementary angle of α, the v sun represents the sun vector, the v cam Denotes the star sensor optical axis vector from the star.

[0024] In this embodiment, the orientation of the optical axis 1 of the star sensor satisfying the sun avoidance constraint in space is determined by the thrust vector arbitrarily oriented from the space of the star.

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Abstract

The invention discloses a decoupling control method for relative orbits and attitudes of formation satellites, relates to the technical field of the control of the orbits and attitudes of a spacecraft formation, and solves the problems of large satellite calculation amount and low orbit solving efficiency caused by high control dimension of the formation satellites due to serious coupling of the relative orbits and the attitudes of the formation satellites. The method gives two decoupling conditions at first, so that the control of the relative orbits and the attitudes can be designed independently; a thrust vector mobility decoupling constraint condition is introduced to the initialization control of the relative orbits according to satellite attitude mobility constraints indirectly; and during the optimal thrust vector attitude tracking, the possible orientation in space, which meets solar avoidance constraints, of a star sensor optical axis (1) is sought by using a geometric method, and the optimal attitude quaternion and attitude angular velocity are calculated finally by using a double-vector attitude determination algorithm. The method provides important reference value for the control of the orbits and the attitudes of the spacecraft formation.

Description

technical field [0001] The invention relates to the technical field of spacecraft formation orbit and attitude control. Background technique [0002] Formation satellites are usually microsatellites, which are small in size and relatively large in payload, so that the quality and space allocated to the propulsion system are often limited to a certain extent. Therefore, it is considered that each satellite in the formation satellites is only equipped with a single continuous small-thrust thruster; At the same time, in order to complete specific space science tasks, in the process of formation initialization, formation reconstruction and formation maintenance, there are high requirements for satellite attitude determination and control accuracy, so other sensors such as star sensors are considered The resulting attitude constraints, for example, the optical axis of the star sensor should not be aligned with the sun during the formation control process. However, due to the con...

Claims

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Application Information

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IPC IPC(8): G05D1/08G05D1/10
Inventor 曹喜滨吴云华张锦绣张世杰
Owner HARBIN INST OF TECH
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