本文關(guān)鍵詞：快速機(jī)動衛(wèi)星姿態(tài)確定與控制算法研究　出處：《哈爾濱工業(yè)大學(xué)》2016年碩士論文　論文類型：學(xué)位論文

  更多相關(guān)文章： 姿態(tài)確定 非線性濾波 姿態(tài)控制 時(shí)間最優(yōu)路徑

【摘要】：航天器的大角度快速機(jī)動中的姿態(tài)確定問題已經(jīng)成為近年來的研究熱點(diǎn)與難點(diǎn),高精度、大角度的快速機(jī)動任務(wù)對衛(wèi)星姿態(tài)確定系統(tǒng)與姿態(tài)控制系統(tǒng)同時(shí)提出了較高的精度要求,針對這一背景,本次研究主要圍繞快速機(jī)動衛(wèi)星姿態(tài)確定與控制問題的以下方面進(jìn)行:快速機(jī)動衛(wèi)星姿態(tài)確定算法的設(shè)計(jì)。首先,針對姿態(tài)動力學(xué)模型非線性強(qiáng)的特性,引入誤差模型的概念,基于最小模型誤差原理,給出了模型誤差的狀態(tài)估計(jì)方程,對于濾波系統(tǒng)模型進(jìn)行修正,從而實(shí)現(xiàn)對于濾波精度的提升;針對星敏感器精度高但采樣周期長,陀螺儀精度相對較低但采樣周期短的特性,引入互補(bǔ)濾波算法,給出了對于衛(wèi)星姿態(tài)確定系統(tǒng)的互補(bǔ)濾波算法;針對近似線性模型與真實(shí)模型相差較大的特性,引入魯棒濾波算法,將系統(tǒng)誤差視為系統(tǒng)模型中的不確定項(xiàng),通過矩陣不等式變換將誤差方差矩陣限定在某一可接受范圍內(nèi),從而實(shí)現(xiàn)對于系統(tǒng)狀態(tài)的估計(jì)。其中基于最小模型誤差原理的改進(jìn)EKF算法與基于互補(bǔ)濾波算法的姿態(tài)確定算法均屬于本次研究的創(chuàng)新內(nèi)容。快速機(jī)動衛(wèi)星時(shí)間最優(yōu)路徑的設(shè)計(jì)。采用偽譜法與配點(diǎn)法對于衛(wèi)星姿態(tài)機(jī)動進(jìn)行路徑規(guī)劃,基于輸出力矩處于飽和狀態(tài)時(shí)有時(shí)間最優(yōu)路徑的假設(shè)條件,給出了衛(wèi)星姿態(tài)機(jī)動的時(shí)間最優(yōu)路徑。對于衛(wèi)星姿態(tài)機(jī)動的時(shí)間最優(yōu)問題進(jìn)行描述,引入配點(diǎn)法與偽譜法,給出了基于這兩種方法的非線性規(guī)劃模型,對于衛(wèi)星姿態(tài)機(jī)動中的三軸耦合與三軸獨(dú)立情形分別進(jìn)行分析,給出了rest to rest機(jī)動的時(shí)間最優(yōu)路徑。同時(shí)基于傳統(tǒng)的Bang-Bang控制律設(shè)計(jì)衛(wèi)星單軸機(jī)動的最優(yōu)路徑,對二者的結(jié)果進(jìn)行了對比�？焖贆C(jī)動衛(wèi)星姿態(tài)跟蹤控制律的設(shè)計(jì)。對于前文提出的姿態(tài)時(shí)間最優(yōu)路徑設(shè)計(jì)跟蹤控制律,引入有限時(shí)間控制,設(shè)計(jì)了基于積分滑模與終端函數(shù)的有限時(shí)間姿態(tài)跟蹤控制律,以實(shí)現(xiàn)衛(wèi)星真實(shí)軌跡在有限時(shí)間內(nèi)追蹤上期望軌跡。同時(shí)考慮到衛(wèi)星不可能一直處于快速機(jī)動的狀態(tài),并考慮到輸出力矩上限,設(shè)計(jì)了有限輸出力矩跟蹤控制律。
[Abstract]:The problem has become a research hotspot and difficulty in recent years, the large angle attitude maneuver of spacecraft fast high precision, fast maneuver with large angle and put forward higher requirements on the accuracy of satellite attitude determination system and attitude control system, based on this background, this research mainly focuses on the following aspects of rapid maneuver of satellite attitude determination with the problem of control: fast maneuvering satellite attitude determination algorithm design. Firstly, according to the characteristics of strong nonlinear attitude dynamics model, introducing the concept of error model, based on the principle of minimum model error, model error estimation equation is given, for filtering system model was modified, in order to achieve the filtering accuracy improvement based on star; sensor with high precision but the long sampling period, the gyroscope precision is relatively low but the short sampling cycle characteristics, the introduction of complementary filtering algorithm for For the satellite attitude determination system complementary filtering algorithm; the approximate linear model and real model with different characteristics, a robust filtering algorithm, the system error as the system model uncertainties, the error variance matrix transform matrix inequality will limit in an acceptable range, so as to realize the estimation of the the state of the system. The improved EKF algorithm based on the principle of minimum model error and complementary filtering algorithm of attitude determination based on the content of the innovation of algorithm belong to this study. Fast maneuvering satellite time optimal path design. And using the method of pseudo spectral collocation method for satellite attitude maneuver path planning, assumption of output torque at saturation when a time optimal path based on a given satellite attitude maneuver time optimal path for time optimal problem of satellite attitude maneuver Description, introduction of collocation method and pseudospectral method, a nonlinear programming model on the basis of these two methods are given for three axis coupling satellite attitude maneuver in the case of independent and three axis were analyzed, given the rest to rest mobile time optimal path. At the same time optimal path motor single axis traditional Bang-Bang law design satellite control based on the comparison of the results of the two. Fast maneuvering satellite attitude tracking control law is designed for the attitude. The proposed design time optimal path tracking control law, the introduction of the finite time control, tracking control law is designed in finite time attitude and terminal function based on integral sliding mode, in order to achieve the true trajectory in Satellite Co. time tracking the desired trajectory. At the same time taking into account the satellite could not have been in rapid maneuver, and considering the output torque limit, the output torque tracking design Co. Control law.

【學(xué)位授予單位】：哈爾濱工業(yè)大學(xué)
【學(xué)位級別】：碩士
【學(xué)位授予年份】：2016
【分類號】：V448.2

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