space sciences and exploration
Ebrahim Amiri; Masoome Khani Chamani; Mahdi Jafari-Nodoshan; Sajjad Ghazanfarinia; Masoud Khoshsima
Articles in Press, Accepted Manuscript, Available Online from 16 December 2023
Abstract
The economic model generally expresses the mechanisms used to earn money from a business, and if it doesnot generate income, its failure will be certain. Therefore, the decision to carry out a mission isnot only based on technical specifications, and besides that, economic profitability is another part ...
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The economic model generally expresses the mechanisms used to earn money from a business, and if it doesnot generate income, its failure will be certain. Therefore, the decision to carry out a mission isnot only based on technical specifications, and besides that, economic profitability is another part of decision-making and will be one of the main factors for commercial investments. Space projects, as wellas moon-mining projects, are no exception to this rule and require an all-round approach to compare financial and technical feasibility. Analyzing the economic feasibility of any project can be summed up in the evaluation of its economic model. In this regard, a model is needed to compare, rank and determine the available options, which is economically justified. In this paper, the economic evaluation of moon-mining based on the materials available on the moon and sending them to the earth is discussed. Materials with economic priority are categorized and selected in a fuzzy evaluation and using an economic model suitable for space mining, an economic evaluation for the business of selling materials on the Earth is carried out and according to economic efficiency, the type of material and also the high-level specifications of the project has been extracted.
Space subsystems design: (navigation, control, structure and…)
Mahdi Jafari Nadoushan; Kosar Aramkhah
Volume 14, Issue 1 , March 2021, , Pages 1-13
Abstract
In this paper, the dumbbell model is used for gravity field of asteroid 216 Kleopatra. Utilizing the model results in governing equations of motion of a spacecraft around an asteroid similar to those of motion of a spacecraft in the restricted circular three-body problem. The equilibrium points and Jacobi ...
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In this paper, the dumbbell model is used for gravity field of asteroid 216 Kleopatra. Utilizing the model results in governing equations of motion of a spacecraft around an asteroid similar to those of motion of a spacecraft in the restricted circular three-body problem. The equilibrium points and Jacobi regions are calculated and symmetric periodic orbits are computed utilizing grid search and shooting methods. The xz-plane is considered as the symmetry plane. Stability of the periodic orbits is evaluated by Floquet theory that shows all the computed orbits are unstable. By adding the solar radiation pressure term to the governing equations of motion, the symmetric periodic orbits are recomputed and index of stability are examined. The results show that the solar radiation pressure, though change the values of the index of stability, does not affect the stability of computed periodic orbits. Therefore, stabilizing a spacecraft on the unstable periodic orbits requires controlling spacecraft.
Space subsystems design: (navigation, control, structure and…)
Maziar Shefaee Roshan; Mahdi Ghobadi; Mahdi jafari Nadoushan
Volume 13, Issue 1 , March 2020, , Pages 71-82
Abstract
Using linear programming method in control allocation for attitude control subsystem of spacecraft with redundant thrusters is studied in this paper. The simplex algorithm is utilized as a solver and the Direction Preserving and Bodson’s Reduced size Direction preserving approaches are used as ...
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Using linear programming method in control allocation for attitude control subsystem of spacecraft with redundant thrusters is studied in this paper. The simplex algorithm is utilized as a solver and the Direction Preserving and Bodson’s Reduced size Direction preserving approaches are used as optimal approaches to deal with non admissible solutions. Also, proper functioning of these approaches against thrusters fault phenomenon is evaluated. The results show that the Direction Preserving approache has less computational time and less fuel consuming. However, the Bodson’s Reduced size Direction preserving approache has more computational time and more fuel consuming but less tri-axis tracking error. It should be noted that the PD controller has been used as a spacecraft control rule, and simulations have been made for the number and configuration of the specific thrusters.
Hassan Naseh; Mehran Mirshams; Elyas Fadakar; Mehdi Jafari Nadoushan
Volume 11, Issue 2 , September 2018, , Pages 47-53
Abstract
The main goal of this paper is to introduce the Moon exploration mission design based on existing technology.The Moon exploration mission design entailsoptimal maneuvering orbit, payload and launch vehicle design. Optimal maneuvering orbit is designed with respect to Circular Restricted Three Body Problem ...
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The main goal of this paper is to introduce the Moon exploration mission design based on existing technology.The Moon exploration mission design entailsoptimal maneuvering orbit, payload and launch vehicle design. Optimal maneuvering orbit is designed with respect to Circular Restricted Three Body Problem (CRTBP) to model the motion of a spacecraft in the Earth/Moon system. To this end, optimal maneuvering orbitadopted CRTBP as dynamical model and obtained three-dimensional Earth to Moon transfers with low cost. This method is more preferable and flexible than Hohmann transfer because of its lower cost and its access to various inclinations in departure and arrival.The optimal Launch Vehicle Conceptual Design (LVCD) algorithm is based on optimization of major design parameters. LVCD algorithm is coded in a software to let the design engineer explore the design space and to reduce the cost and time of the conceptual design phase that is developed by the authors.The optimization process is performed subject to the restrictions and the performance index is optimized in a mutual iteration mechanism. Consequently, the designed launch vehicle ability to satisfy the mission objectives and its requirements is evaluated.
M. Jafari -Nadoushan; A. Novinzadeh
Volume 6, Issue 3 , October 2013, , Pages 49-54
Abstract
In this paper design of transfer trajectory from Earth park orbit to a halo orbit around L1 of Earth-Moon system and return trajectory from halo orbit to the Earth are investigated. Since satisfying constraints and boundary conditions at the end of trajectory is an important point in trajectory design, ...
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In this paper design of transfer trajectory from Earth park orbit to a halo orbit around L1 of Earth-Moon system and return trajectory from halo orbit to the Earth are investigated. Since satisfying constraints and boundary conditions at the end of trajectory is an important point in trajectory design, we deal with a two point boundary value problem. Considered constraints in this paper include height, orthogonality of position and velocity vectors for reducing required Del-V for orbital transfer and flight path angle. Due to complex dynamics of three body problem and also in order to satisfying these constraints and suitable trajectory design, the multiple shooting methods based on differential correction is used.
M. Jafari Nadoushan; M. Tivay
Volume 3, Issue 2 , January 2011, , Pages 53-58
Abstract
Effect of regularization on the solution of perturbed two body problem is investigated in this paper. Purposes of using this method are computational burden reduction and achieving desirable accuracy in the minimum time. In this regard the equations of motion are linearized and independent variable is ...
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Effect of regularization on the solution of perturbed two body problem is investigated in this paper. Purposes of using this method are computational burden reduction and achieving desirable accuracy in the minimum time. In this regard the equations of motion are linearized and independent variable is changed from time to the true anomaly. These yield reducing run time, however increasing accuracy. The results of simulation confirm that utilizing this method in onboard computation or long term simulations is more suitable and efficient than other methods including general and special perturbation methods.
M. Jafari Nadoushan; S. H. Pourtakdoust
Volume 3, Issue 1 , July 2010, , Pages 75-80
Abstract
Development of halo orbits and their associated invariant manifolds are investigated. Halo orbits play a fundamental role in complex space mission designs. In essence, halo orbits are periodic solutions of the restricted three body problem (R3BP) determined under specific initial conditions. In this ...
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Development of halo orbits and their associated invariant manifolds are investigated. Halo orbits play a fundamental role in complex space mission designs. In essence, halo orbits are periodic solutions of the restricted three body problem (R3BP) determined under specific initial conditions. In this paper, the symmetric property of the nonlinear R3BP governing differential equations is utilized in order to obtain the desired initial conditions. In this regard the differential correction technique and the state transition matrix are used to generate the halo orbits. The differential correction technique, based on the Newton method, is an effective tool for solving two point boundary value problems. In addition to generate the stable and unstable manifolds, the initial conditions are perturbed in the direction of Eigenvectors and the equations of motion are integrated for an arbitrary time interval.
