نشریه علوم و فناوری فضایی
پیاپی 2 (زمستان 1387)

Solar panels are the primary sources of power in a satellite. Operating characteristics of the solar cells, such as current, voltage and generated power, depend on their operating temperatures and the amount of solar radiation received by the solar cells. Therefore, for optimum operation of the solar cells, it is essential to control their temperatures within acceptable limits, and provide the maximum possible solar radiation for the solar cells. Solar panel configurations include fixed and deployable panels; the latter configuration being flexible, providing the possibility of sun tracking for maximum utilization of solar radiation. In this paper we have considered a cubic satellite, having four deployable solar panels on its lateral sides, which can be deployed at certain angle (called deployment angle) with respect to the satellite body. Four limiting values of beta angle (angle between solar vector and orbital plane) have been considered, and for each beta angle, various solar panel deployment angles have been studied. The amounts of radiations received by the cells for each deployment angle have been presented. The solar panels have been modeled and thermally analyzed, to determine temperatures of the solar cells at various beta angles, and for different panel deployment angles. Results show that for the beta angles considered, and the satellite under study, a 30° solar panel deployment angle presents the optimum conditions for the operation of the solar cells.

Attitude control system of satellite with Gravity Gradient stabilization requires high moments of inertia ratio for providing stability and continuous orientation toward Earth. Although, this high ratio causes satellite has small body and reduce mission capability. In this paper, moments of inertia ratio is reduced using a closed form formula based on our previous work, in such a way that it could be provided more missions by augmented solar panels to satellite. Solar orientation could be yielded by rotating satellite about gravity gradient boom (yaw rotation). Interaction between yaw rotation and satellite rotation around Earth causes biased-attitude error in roll axis. To overcome this problem, it is necessary to reduce yaw rotation by adding a reaction wheel toward boom direction. To realization this method, stability criteria of gravity gradient is developed and control law for small and large angles rotation is designed in such a way that angular momentum and moment constraints of reaction wheel to be satisfied. Finally, fine performance of attitude control system will be illustrated with simulation based on specification of an on-orbit satellite and actual consideration.

The principle goal of this paper is to introduce Launch Vehicle Conceptual Design (LVCD) software based on multi-parameter optimization idea. The main objectives of this software arereduction of the cost and time of conceptual design phase. This software is user friendly such that an operator familiar with fundamentals of design and launch vehicle mass – energy equations and with primary training operator is capable to work with LVCD.The algorithm used in LVCD, is based on combinational optimization of major design parameters. To this end, ten sub-algorithms will be presented in this design approach. Mass distribution of different stages to launch maximum payload mass to the orbit, pitch program trajectory to get to the maximum final velocity, and providing minimum velocity loss due to gravity, and also minimum axial acceleration of various stages of launch vehicle will be optimized as the results of the presented approach. The optimization process is performed subject to the restrictions. Also, the performance index is optimized in a mutual iteration mechanism (multi-parameter optimization). Evaluation and verification of the presented method is performed using available data of two and three-stage launch vehicles

Attitude control system of satellite with Gravity Gradient stabilization requires high moments of inertia ratio for providing stability and continuous orientation toward Earth. Although, this high ratio causes satellite has small body and reduce mission capability. In this paper, moments of inertia ratio is reduced using a closed form formula based on our previous work, in such a way that it could be provided more missions by augmented solar panels to satellite. Solar orientation could be yielded by rotating satellite about gravity gradient boom (yaw rotation). Interaction between yaw rotation and satellite rotation around Earth causes biased-attitude error in roll axis. To overcome this problem, it is necessary to reduce yaw rotation by adding a reaction wheel toward boom direction. To realization this method, stability criteria of gravity gradient is developed and control law for small and large angles rotation is designed in such a way that angular momentum and moment constraints of reaction wheel to be satisfied. Finally, fine performance of attitude control system will be illustrated with simulation based on specification of an on-orbit satellite and actual consideration

In this paper, a new approach is proposed for solving the problem of optimal low thrust orbit transfer. In this approach, the problem of trajectory optimization of optimal orbit transfer is defined by modified equinoctial orbital elements. For solving this problem, direct collocation method, that is an efficient numerical method for solving optimal control problems, is used. By using this method, the problem of trajectory optimization is fully discretized and converted to a nonlinear programming problem. This discrete problem with large numbers of variables and constraints is solved by a powerful nonlinear programming solver (IPOPT). Finally, optimal state and control variables are achieved for optimal orbit transfer with minimum fuel consumption.

In this study, optimal low-thrust spacecraft trajectories are obtained by time-domain finite element method. Equations of motion are expressed in state-space form. The performance index is considered as minimum time. The problem has been formulated through the variational approach. The time-domain finite element discretized form of the performance index, state equation constraints and the related boundary conditions are presented. By setting out the discrete equations, a set of nonlinear algebraic equations is generated and by using Newton–Raphson method, optimum answer is attained. The effects of the number of time segments on the performance index are examined. Furthermore, the influences of effective exhaust velocities on the optimal trajectory are demonstrated.

In this paper, based on the Lagrange method, attitude motion equations for a flexible spacecraft have been derived. Flexible appendages are modeled by Euler-Bernoulli beam. Hybrid control scheme of H∞ and sliding mode are used for attitude regulation. Switching between these to algorithm is determined using absolute error parameter, so that when this parameter is large (i.e. the attitude is far from its desired conditions) the sliding mode control is used. in contrast, when the spacecraft is close to the desired attitude, the control is based on H∞ method. This hybrid scheme leads to a fast response and also robustness against uncertainty. Switching surface has been designed so that a certain cost function is minimized. In H∞ controller design, the first three vibration modes of the flexible spacecraft are considered as well as the Euler angles and their rates.

This paper provides an instruction for designing and building of an Inverted F antenna mounted on a cylindrical conducting body with a conical nose. Designed antennas were simulated using a full wave simulator HFSS based on the finite element method and their radiation patterns and return loss were studied. The simulated data were compared with measurement results.