The Experts below are selected from a list of 40122 Experts worldwide ranked by ideXlab platform
Tamer Mekky Ahmed Habib - One of the best experts on this subject based on the ideXlab platform.
-
Combined Spacecraft Orbit and attitude control through extended Kalman filtering of magnetometer, gyro, and GPS measurements
The Egyptian Journal of Remote Sensing and Space Science, 2014Co-Authors: Tamer Mekky Ahmed HabibAbstract:Abstract The main goal of this research is to establish Spacecraft Orbit and attitude control algorithms based on extended Kalman filter which provides estimates of Spacecraft Orbital and attitude states. The control and estimation algorithms must be capable of dealing with the Spacecraft conditions during the detumbling and attitude acquisition modes of operation. These conditions are characterized by nonlinearities represented by large initial attitude angles, large initial angular velocities, large initial attitude estimation error, and large initial position estimation error. All of the developed estimation and control algorithms are suitable for application to the next Egyptian scientific satellite, EGYPTSAT-2. The parameters of the case-study Spacecraft are similar but not identical to the former Egyptian satellite EGYPTSAT-1. This is done because the parameters of EGYPTSAT-2 satellite have not been consolidated yet. The sensors utilized are gyro, magnetometer, and GPS. Gyro and magnetometer are utilized to provide measurements for the estimates of Spacecraft attitude state vector where as magnetometer and GPS are utilized to provide measurements for the estimates of Spacecraft Orbital state vector.
-
Simultaneous Spacecraft Orbit estimation and control based on GPS measurements via extended Kalman filter
The Egyptian Journal of Remote Sensing and Space Science, 2013Co-Authors: Tamer Mekky Ahmed HabibAbstract:The primary aim of this work is to provide simultaneous Spacecraft Orbit estimation and control based on the global positioning system (GPS) measurements suitable for application to the next coming Egyptian remote sensing satellites. Disturbance resulting from earth’s oblateness till the fourth order (i.e., J4) is considered. In addition, aerodynamic drag and random disturbance effects are taken into consideration.
-
Global optimum Spacecraft Orbit control subject to bounded thrust in presence of nonlinear and random disturbances in a low earth Orbit
The Egyptian Journal of Remote Sensing and Space Science, 2012Co-Authors: Tamer Mekky Ahmed HabibAbstract:Abstract The primary objective of this work is to develop an effective Spacecraft Orbit control algorithm suitable for Spacecraft Orbital maneuver and/or rendezvous. The actual governing equation of a Spacecraft Orbiting the earth is merely nonlinear. Disturbance forces resulting from aerodynamic drag, oblateness of the earth till the fourth order (i.e. J 4 ), and random disturbances are modeled for the initial and target Orbits. These disturbances increase the complexity of nonlinear governing equations. Global optimum solutions of the control algorithm parameters are determined throughout real coded genetic algorithms such that the steady state difference between the actual and desired trajectories is minimized. The resulting solutions are constrained to avoid Spacecraft collision with the surface of the earth taking into account limited thrust budget.
Xingqun Zhan - One of the best experts on this subject based on the ideXlab platform.
-
Spacecraft Orbit propagator integration with gnss in a simulated scenario
Advances in Space Research, 2017Co-Authors: Shuai Jing, Xingqun ZhanAbstract:Abstract When space vehicles operate above the Global Navigation Satellite System (GNSS) constellation or even above geosynchronous Orbit, it is common that the traditional GNSS single–epoch solution can’t meet the requirement of Orbit determination (OD). To provide the required OD accuracy continuously, a new designed Spacecraft Orbit propagator (OP) is combined with the GNSS observations in a deep integration mode. Taking both the computational complexity and positioning accuracy into consideration, the Orbit propagator is optimized based on a simplified fourth order Runge-Kutta integral aided with empirical acceleration model. A simulation scenario containing a typical Highly-inclined Elliptical Orbit (HEO) user and GPS constellation is established on a HwaCreat™ GNSS signal simulator to testify the performance of the design. The numerical test results show that the maximum propagation error of the optimized Orbit propagator does not exceed 1000 m within a day, which is superior to conventional OPs. If the new OP is deeply integrated with GNSS in our proposed scheme, the 95% SEP for the OD accuracy is 10.0005 m, and the time to first fix (TTFF) values under cold and warm start conditions are reduced by at least 7 s and 2 s respectively, which proves its advantage over loose integration and tight integration.
Yuriy Nikolaevich Chelnokov - One of the best experts on this subject based on the ideXlab platform.
-
Pulsed Optimal Spacecraft Orbit Reorientation by Means of Reactive Thrust Orthogonal to the Osculating Orbit. II
Mechanics of Solids, 2019Co-Authors: Yakov Grigorievich Sapunkov, Yuriy Nikolaevich ChelnokovAbstract:A new theory and a new algorithm for numerical solution of the problem of optimal Spacecraft Orbit reorientation by means of a pulsed (large) thrust, orthogonal to the osculating Orbit plane, for a non-fixed number of thrust impulses are set out in a strict non-linear formulation. As a control, a vector of reactive acceleration from engine thrust is used. The combined functional is minimized, equal to the weighted sum of the reorientation time and the reactive acceleration impulse (characteristic speed) during the time of the Spacecraft Orbit reorientation (a special case of this functional is the case of minimization of the characteristic speed). To construct the theory, a solution for the problem of optimal Spacecraft Orbit reorientation in continuous formulation (using limited (low) thrust) is described in the first part of this article. It is shown that the problem of optimal impulse Spacecraft Orbit reorientation in the case when optimal control consists of two reactive acceleration impulses, applied to the Spacecraft at initial and final moments of time of motion, is solved analytically. Examples of numerical solution of the problem of optimal impulse Spacecraft Orbit reorientation are given, illustrating the capabilities of the proposed method.
-
Pulsed Optimal Spacecraft Orbit Reorientation by Means of Reactive Thrust Orthogonal to the Osculating Orbit. I
Mechanics of Solids, 2018Co-Authors: Yakov Grigorievich Sapunkov, Yuriy Nikolaevich ChelnokovAbstract:The first part of the article provides an overview of the work on the differential equations of the Spacecraft (SC) Orbit orientation and the problem of optimal reorientation of a Spacecraft Orbit in an inertial coordinate system by means of reactive acceleration orthogonal to the osculating plane of the Spacecraft. The theory of solving the problem of the optimal reorientation of the Orbit of the Spacecraft using the quaternionic differential equation for the orientation of the Orbital coordinate system in a non-linear continuous formulation (using limited (small) thrust) is presented. As a minimized quality functional, a combined functional is used equal to the weighted sum of the reorientation time and thrust impulse (characteristic speed) during the reorientation of the Orbit of the Spacecraft (special cases of this functional are the speed response case and the characteristic speed minimization separately).
-
Optimal reorientation of Spacecraft Orbit
Archives of Control Sciences, 2014Co-Authors: Yuriy Nikolaevich Chelnokov, Ilya Alekseevich Pankratov, Yakov Grigorievich SapunkovAbstract:Abstract The problem of optimal reorientation of the Spacecraft Orbit is considered. For solving the problem we used quaternion equations of motion written in rotating coordinate system. The use of quaternion variables makes this consideration more efficient. The problem of optimal control is solved on the basis of the maximum principle. An example of numerical solution of the problem is given.
-
The Use of Quaternions in the Optimal Control Problems of Motion of the Center of Mass of a Spacecraft in a Newtonian Gravitational Field: I
Cosmic Research, 2003Co-Authors: Yuriy Nikolaevich ChelnokovAbstract:The problem of optimal control is considered for the motion of the center of mass of a Spacecraft in a central Newtonian gravitational field. For solving the problem, two variants of the equations of motion for the Spacecraft center of mass are used, written in rotating coordinate systems. Both the variants have a quaternion variable among the phase variables. In the first variant this variable characterizes the orientation of an instantaneous Orbit of the Spacecraft and (simultaneously) the Spacecraft location in this Orbit, while in the second variant only the instantaneous Orbit orientation is specified by it. The suggested equations are convenient in the respect that they allow the general three-dimensional problem of optimal control by the motion of the Spacecraft center of mass to be considered as a composition of two interrelated problems. In the first variant these problems are (1) the problem of control of the shape and size of the Spacecraft Orbit and (2) the problem of control of the orientation of a Spacecraft Orbit and the Spacecraft location in this Orbit. The second variant treats (1) the problem of control of the shape and size of the Spacecraft Orbit and the Orbit location of the Spacecraft and (2) the problem of control of the orientation of the Spacecraft Orbit. The use of quaternion variables makes this consideration most efficient. The problem of optimal control is solved on the basis of the maximum principle. Several first integrals of the systems of equations of the boundary value problems of the maximum principle are found. Transformations are suggested that reduce the dimensions of the systems of differential equations of boundary value problems (without complicating them). Geometrical interpretations are given to the transformations and first integrals. The relation of the vectorial first integral of one of the derived systems of equations (which is an analog of the well-known vectorial first integral of the studied problem of optimal control) with the found quaternion first integral is considered. In this paper, which is the first part of the work, we consider the models of motion of the Spacecraft center of mass that employ quaternion variables. The problem of optimal control by the motion of the Spacecraft center of mass is investigated on the basis of the first variant of equations of motion. An example of a numerical solution of the problem is given.
Han Chao - One of the best experts on this subject based on the ideXlab platform.
-
Development of Simulation Software for High Precision Spacecraft Orbit Prediction
Computer Simulation, 2008Co-Authors: Han ChaoAbstract:Spacecraft Orbit prediction is very necessary in the process of Spacecraft design,and high precision or- bit prediction simulation software is a very important tool for Spacecraft Orbit prediction.In this paper the theory of Spacecraft Orbit prediction was briefly introduced.The method for modeling Spacecraft Orbit dynamics was presented in detail.The idea of designing high precision Spacecraft Orbit prediction simulation software was studied,and how to re- alize such software was given.At last,a software called"Spacecraft Orbit Calculation Tool(SOCT)"was introduced and tested by Satellite Tool Kit(STK),which proved that SOCT has achieved the standard of high precision Orbit pre- diction.
-
Realization of A Small System of Spacecraft Orbit Design
Computer Engineering, 2001Co-Authors: Han ChaoAbstract:This paper introduces the functions of the Spacecraft Orbit design system briefly, then describes the realization scheme, the modules and the structure of the small system detailedly. At last, it points out the kernel technology and summarizes the performance of the system.
Shuai Jing - One of the best experts on this subject based on the ideXlab platform.
-
Spacecraft Orbit propagator integration with gnss in a simulated scenario
Advances in Space Research, 2017Co-Authors: Shuai Jing, Xingqun ZhanAbstract:Abstract When space vehicles operate above the Global Navigation Satellite System (GNSS) constellation or even above geosynchronous Orbit, it is common that the traditional GNSS single–epoch solution can’t meet the requirement of Orbit determination (OD). To provide the required OD accuracy continuously, a new designed Spacecraft Orbit propagator (OP) is combined with the GNSS observations in a deep integration mode. Taking both the computational complexity and positioning accuracy into consideration, the Orbit propagator is optimized based on a simplified fourth order Runge-Kutta integral aided with empirical acceleration model. A simulation scenario containing a typical Highly-inclined Elliptical Orbit (HEO) user and GPS constellation is established on a HwaCreat™ GNSS signal simulator to testify the performance of the design. The numerical test results show that the maximum propagation error of the optimized Orbit propagator does not exceed 1000 m within a day, which is superior to conventional OPs. If the new OP is deeply integrated with GNSS in our proposed scheme, the 95% SEP for the OD accuracy is 10.0005 m, and the time to first fix (TTFF) values under cold and warm start conditions are reduced by at least 7 s and 2 s respectively, which proves its advantage over loose integration and tight integration.