Space subsystems design: (navigation, control, structure and…)
Mahdi Rivandi; Mehran Mirshams; Mohammad Zarourati
Volume 16, Issue 1 , March 2023, , Pages 75-88
Abstract
To test the Attitude Determination and Control Subsystem of a satellite, it is necessary to have an attitude dynamics simulator, and the simulator must be in a balance condition. Disturbances on the balance system in the simulation include deviations caused by the difference between the center of mass ...
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To test the Attitude Determination and Control Subsystem of a satellite, it is necessary to have an attitude dynamics simulator, and the simulator must be in a balance condition. Disturbances on the balance system in the simulation include deviations caused by the difference between the center of mass and rotation, as well as the movement of two horizontal actuators. The movement of two horizontal actuators is a factor for rotational and vortex motion. In the simulation of experimental models, PID control coefficients are also used to control three axes. The balance system actuators include moving masses and reaction wheel that are installed around the horizontal and vertical axes, respectively. To validate the results, a hardware sample has been developed for laboratory tests. Using the sampling time, models and experimental coefficients, the hardware reaches the accuracy of 0.2 and 0.5 degrees in 25 seconds, respectively, which indicates a suitable accuracy for balancing the simulator of the CubeSat attitude.
GPS and navigation GPS)، GLONASS، GALILEO
Mohsen Shamirzaei; Mehran Mir Shams
Volume 14, Issue 3 , September 2021, , Pages 75-90
Abstract
The main task of the study is to estimate the position error in an inertial navigation system by integrating it with the visual system. The case study is a spacecraft that must accurately measure its position relative to a predetermined landing point. The spacecraft is assumed to be augmented GNSS navigation. ...
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The main task of the study is to estimate the position error in an inertial navigation system by integrating it with the visual system. The case study is a spacecraft that must accurately measure its position relative to a predetermined landing point. The spacecraft is assumed to be augmented GNSS navigation. Therefore, when satellite signals are dropped out or when landing on a moving marine platform, the data of the vision navigation system replaces the information of the satellite navigation system and improves the accuracy of the spacecraft navigation system. An Extended Kalman filter has been used to integrate inertial and vision navigation system information. In addition, the output data of the vision system, in order to be used in the Kalman filter measurement equations, is first processed by the recursive least square filter. The relevant relations are given and based on the results of software simulation, the efficiency of the proposed method is shown.
Space subsystems design: (navigation, control, structure and…)
Niki Sajjad; Mehran Mirshams; Shahrokh Jaliian
Volume 13, Issue 3 , September 2020, , Pages 51-62
Abstract
This paper presents design, analysis and performance verification test of student microsatellite Attitude Determination and Control Subsystem (ADCS) . ADCS design and implementation procedure contains several various steps. There are four main test levels during simulation called “Model-in-the-Loop”, ...
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This paper presents design, analysis and performance verification test of student microsatellite Attitude Determination and Control Subsystem (ADCS) . ADCS design and implementation procedure contains several various steps. There are four main test levels during simulation called “Model-in-the-Loop”, “Software-in-the-Loop”, “Processor-in-the-Loop” and “Hardware-in-the-Loop”. This paper is a result of scientific and practical research during two years, on the student microsatellite, which is an eight-nation collaboration project among Asia-Pacific universities. In what follows, “Model-in-the-Loop” and “Processor-in-the-Loop” test and simulation will be discussed. The aim of this paper is to illustrate the result of these two tests and validate the ADCS design. In the end, it is demonstrated that designed control algorithms are precise enough to meet the student microsatellite ADCS requirements and they can be used in the next level of microsatellite development.
mehran mirshams; Mohammad Teshneh lab; Morteza Ramezani
Volume 11, Issue 3 , December 2018, , Pages 1-8
Abstract
Modeling and analyzing systems, especially in complex systems with high dynamics, noise and uncertainty in understanding the behavior of systems and decision making is very important problem from long time ago. This paper shows that neuro-fuzzy systems can be used effectively to design the solar arrays ...
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Modeling and analyzing systems, especially in complex systems with high dynamics, noise and uncertainty in understanding the behavior of systems and decision making is very important problem from long time ago. This paper shows that neuro-fuzzy systems can be used effectively to design the solar arrays of electrical power subsystem of a remote sensing satellite in conceptual design phase. In the design of neuro-fuzzy system, Takagi-Sugeno inference system, hybrid training algorithm and Gaussian membership functions are used. The simulation results obtained in this modeling have an accurate accuracy compared to the experimental data and classical calculations of remote sensing satellites.
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.
Zeynab Aghajani; Ehsan Zabihian; Mehran Mirshams
Volume 10, Issue 4 , March 2018, , Pages 41-54
Abstract
The significance and the wide use of geostationary communication satellites and the long hours of work in the process of their conceptual design was the main motivation to develop a software based on the statistical design to reduce the time spent on the conceptual design phase. This software is based ...
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The significance and the wide use of geostationary communication satellites and the long hours of work in the process of their conceptual design was the main motivation to develop a software based on the statistical design to reduce the time spent on the conceptual design phase. This software is based on the statistical and parametric design method. The statistical model used in this software includes a database of 147 satellites launched between 2010 and 2016. To increase the accuracy of the software, the combined parametric model has been used from selected design references. The software is based on MATLAB and to make it more user friendly, the graphical GUI was used. In this article, the design of the software is presented and there is focus on the design and verification method. The accuracy of this tool was amply verified through a flight prototype, indicating the average error of 16% in the obtained results.
Mehran Mirshams; Ehsan Zabihian
Volume 10, Issue 3 , December 2017, , Pages 1-14
Abstract
This study introduces a new computer code termed AZMIN developed by Space Research Laboratory (SRL). This efficient tool which benefits from the Statistical Design Model (SDM) has been developed for the system design of GEO communication satellites. The main advantage of the AZMIN is to determine technical ...
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This study introduces a new computer code termed AZMIN developed by Space Research Laboratory (SRL). This efficient tool which benefits from the Statistical Design Model (SDM) has been developed for the system design of GEO communication satellites. The main advantage of the AZMIN is to determine technical specification parameters of a satellite at both system and subsystem levels, with a high accuracy and time performance. System-level parameters encompass mass, power, dimension and cost; while, subsystem parameters contain mass, power, and solutions for components configurations of each subsystem. Actual computations of this tool are carried out by means of SDM, leading to a dramatic decrease in the conceptual design time and consequently, its cost. The database utilized is composed of records of 462 GEO communication satellites launched from the year 2000 to 2017. The accuracy of the AZMIN code is amply verified through an example and also a statistical method, demonstrating the mean error of approximately 15% in the obtained results.
Hasan Naseh; Mehran Mirshams; Javad Naderifar
Volume 9, Issue 3 , December 2016, , Pages 73-79
Abstract
The main goal of this paper is development of multi-stage Launch Vehicle (LV) system design software based on advanced classical method. This software has been named Launch Vehicle Conceptual Classical Design (LVCCD). This software covers the complete syllabuses of LV System Design (LVSD) course. The ...
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The main goal of this paper is development of multi-stage Launch Vehicle (LV) system design software based on advanced classical method. This software has been named Launch Vehicle Conceptual Classical Design (LVCCD). This software covers the complete syllabuses of LV System Design (LVSD) course. The main characteristic of the software development is to step by step training the LVSD. Also it can help the better understand in the course in the best quality and lower time. The algorithm used in the software developed according to the outline of LVSD (major design parameters, LV's mass-energy equations and velocity losses and etc.) and using the multi-stage LV statistical data. Hence, these advantages led to better understanding and conceive. Also LVCCD can improve the qualification of training. Finally, the LVCCD software evaluated and verified with the design software as Launch Vehicle Conceptual Design (LVCD) and PBRM by using existing multi-stage LV.
Hojat Taei; M. Mirshams; M. Ghobadi; M. A. Vahid D.; H. Haghi
Volume 8, Issue 4 , January 2016, , Pages 35-44
Abstract
This article describes the details of a Tri-axial Spacecraft Simulator Testbed (TSST) that has been developed as part of a research program on spacecraft multi-body rotational dynamics and control in Space Research Laboratory (SRL) at K. N. Toosi University of Technology. This dumbbell style simulator ...
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This article describes the details of a Tri-axial Spacecraft Simulator Testbed (TSST) that has been developed as part of a research program on spacecraft multi-body rotational dynamics and control in Space Research Laboratory (SRL) at K. N. Toosi University of Technology. This dumbbell style simulator includes a variety of components: spherical air-bearing, inertial measurement unit (IMU), rechargeable battery, reaction wheels (RW), on-board computer (OBC) and balancing masses. In this paper, an attitude control problem for the spacecraft simulator actuated by three reaction wheels is studied. Under the assumption of uniform gravity and frictionless air-bearing environment, reaction wheels generate control moments about the roll, pitch and yaw axes of the base body. The control objective is to perform attitude commands sent from users with the least power consumption and a high precision. To handle the non-linear model, a Linear Quadratic Ricatti (LQR) controller has been programmed and it efficaciously controlled the computer-modeled simulator for any given slewing maneuver. This control approach has been developed to facilitate the system to accomplish large-angle, three-axis slewing maneuvers using RWs as effective actuators.
Mehrn Mirshams; Asad Saghari; Ehsan Zabihian
Volume 8, Issue 3 , October 2015, , Pages 55-63
Abstract
This paper proposes a supplementary method for conceptual design of satellite electrical power subsystem(EPS). Each of represented methods for satellite electricalpower subsystemconceptual design in different references have some advantages and also disadvantages, besides in each of the methods a determined ...
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This paper proposes a supplementary method for conceptual design of satellite electrical power subsystem(EPS). Each of represented methods for satellite electricalpower subsystemconceptual design in different references have some advantages and also disadvantages, besides in each of the methods a determined part of this subsystem has been in focused. In this research, first advantages and disadvantages of existing approaches for the conceptual design of electrical power subsystemwere reviewed, continued with combining of previous methods, improved relationships and using some of the simulation methods plus the using of statistical databases, a complementary method with more ascendency and less disadvantages in comparison with other approaches was presented. Finally, using a data from a specific satellite and the results of the statistical design, the complementary method has been validated.
H. Fazeli; H. Naseh; M. Mirshams; A.B. Novinzadeh
Volume 7, Issue 3 , October 2014, , Pages 9-21
Abstract
Designing space propulsion systems as one of the important subsystems of the spacecrafts and upper stage space launch systems needs to bypass different and complicated steps. In this article the comprehensive process of designing liquid fuel low-thrust space propulsion systems was illustrated. In the ...
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Designing space propulsion systems as one of the important subsystems of the spacecrafts and upper stage space launch systems needs to bypass different and complicated steps. In this article the comprehensive process of designing liquid fuel low-thrust space propulsion systems was illustrated. In the presented pattern, first of all according to the requirements and mission constraints, the main characteristics of the system were determined and then other characteristics were extracted. Finally, for the evaluation of the presented pattern, a low-thrust space propulsion system was designed based on a special mission and the results were compared with a real model. Comparison between the designed space propulsion system and the real one showed an appropriate accuracy of the presented pattern
A. Saghari; M. Mirshams; A. Jafarsalehi
Volume 7, Issue 2 , July 2014, , Pages 35-47
Abstract
In this article the results of research to achieve a comprehensive code of remote sensing satellite conceptual design is presented. In compiling the code with considering the design philosophy of "better, faster, cheaper" has been attempted in caddition to the use of new technologies and the experiences ...
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In this article the results of research to achieve a comprehensive code of remote sensing satellite conceptual design is presented. In compiling the code with considering the design philosophy of "better, faster, cheaper" has been attempted in caddition to the use of new technologies and the experiences of experts, local constraints such as national launchers limitation also be considered.The main advantage of the proposed code than conventional models, is using accurate simulation methods and newer models in the process of conceptual design of subsystems. view of the practical experience of the past to choose the best design starting point, achieve an operational plan to reduce the risk of costly changes, next steps of design has been achieved.
M. Mirshams; S. Irani; A. M. Akhlaghi; H. Naseh
Volume 5, Issue 2 , July 2012, , Pages 49-57
Abstract
The goal of this paper is presenting a methodology for reliability allocation to launch vehicle subsystems using Analytical Hierarchy Process (AHP) method in conceptual design phase. In this methodology, the goal function is reliability and the main considered criterions are technology, complexity, operational ...
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The goal of this paper is presenting a methodology for reliability allocation to launch vehicle subsystems using Analytical Hierarchy Process (AHP) method in conceptual design phase. In this methodology, the goal function is reliability and the main considered criterions are technology, complexity, operational time of each subsystem and cost. For applying AHP method to launch vehicle subsystems reliability allocation, a Matlab code( for investigating compatibility and determining allocation weight factors by employing Matrix Eigen Vector Method) and a Excel sheet( for forming the comparison matrix) are employed. To this point, by using the outcomes of liquid-propellant launch vehicle conceptual design software (LVCD) which developed by authors, the launch vehicle specifications and operational time of each subsystems is derived and is feed to this methodology as input. The results of applying this method to launch vehicle reliability allocation for the second stage of a launch vehicle, shows the error of this method below 2%. It is clear that this small error in reliability issues in conceptual design phase is acceptable.
M. Mirshams; L. Khalaj-Zade
Volume 4, Issue 2 , January 2012, , Pages 11-22
Abstract
To design a manned spacecraft carrying one to two crews to the low Earth orbits, design phases should be completed in various levels. It also needs to gather manned spacecrafts technical data which is developed in the same category. In the system design algorithm presented in this paper, the conceptual ...
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To design a manned spacecraft carrying one to two crews to the low Earth orbits, design phases should be completed in various levels. It also needs to gather manned spacecrafts technical data which is developed in the same category. In the system design algorithm presented in this paper, the conceptual design sequences of a manned spacecraft named Dousti is accomplished systematically.
First of all, in accordance with a target group of manned spacecrafts’ mission, Dousti’s mission profile is defined and system level requirements are recognized. User’s requirements are also considered in the mission profile and subsequently in system level requirements.
General characteristics of Dousti spacecraft as well as its mass and dimensional features are derived in the next step. Statistics and parametric models are systematically applied in design sequence. Then, final characteristics of the spacecraft’s main subsystems designed through engineering methods and applying parametric models are introduced.
Afterwards, resulting characteristics of the spacecraft are traded off to reform and then validated by statistics and parametric models to present the final plan.
M. Mirshams; L. Khaladjzadeh
Volume 3, Issue 1 , July 2010, , Pages 25-36
Abstract
Designing a manned spacecraft carrying one or two persons to low Earth orbits needs to recognize system level requirements and acquire technical data developed in this eria. Revising manned spacecrafts’ characteristics leads to recognize system level requirements and achieve applicable results ...
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Designing a manned spacecraft carrying one or two persons to low Earth orbits needs to recognize system level requirements and acquire technical data developed in this eria. Revising manned spacecrafts’ characteristics leads to recognize system level requirements and achieve applicable results which are needed to design and development of such a spacecraft. Manned spacecraft characteristics comparing charts and figures show a roughly analogous pattern in terms of mass and dimensions and confirm the parallel subsystems have similar performance.
A. Tavakoli; M. Nikusokhan; J. Roshanian; M. Mirshams
Volume 2, Issue 2 , July 2009, , Pages 51-60
Abstract
Design of launch vehicle (LV) trajectory is among the problems in which the use of optimization is of high significance. Implementing optimization using optimal control problem leads to a two point boundary value problem (TPBVS) that can be solved only numerically. On the other hand, development of optimal ...
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Design of launch vehicle (LV) trajectory is among the problems in which the use of optimization is of high significance. Implementing optimization using optimal control problem leads to a two point boundary value problem (TPBVS) that can be solved only numerically. On the other hand, development of optimal control problem for sophisticated model is very intricate and therefore simulation-based optimization plays an Important role in these problems. In this paper, a LV trajectory defining control input as a parameteic function with linear, Spline and Bezier functions was designed and its fuel consumption was optimized using Genetic Algorithm. Result analyses speculate that Bezier and Spline functions arrives to favorable consequences in terms of meeting terminal Boundary Condition (B. C), optimality of LV payload and also number of optimization parameters.
S.H. Miri Roknabadi; M. Mirshams; A. A. Nikkhah
Volume 2, Issue 2 , July 2009, , Pages 61-68
Abstract
This paper presents a technical note of mathematic model, design and manufacturing steps of a Reaction Wheel, one of the most important active actuators of satellite. After that Reaction Wheels are tested for the satellite simulator of K.N.Toosi University of Technology, Iran. There were some requirements ...
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This paper presents a technical note of mathematic model, design and manufacturing steps of a Reaction Wheel, one of the most important active actuators of satellite. After that Reaction Wheels are tested for the satellite simulator of K.N.Toosi University of Technology, Iran. There were some requirements and restrictions such as needed maximum torque and control accuracy for attitude maneuver, receivable power, voltage and current. Accordingly fundamental components of Reaction Wheel have been designed and selected. Wheel, motor, bearings and retentive are the significant components. At the rest of the paper, the substantial parameters of the Reaction Wheels are confirmed by a new test set. The results of test guarantee a satisfactory stabilization and accurate maneuver.
M. Mirshams; H. Karimi; H. Naseh
Volume 1, Issue 2 , December 2008, , Pages 17-25
Abstract
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 ...
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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.
M. Mirshams; H. Karimi; H. Naseh
Volume 1, Issue 1 , September 2008, , Pages 21-36
Abstract
The principle goal of this paper is developing of Launch Vehicle Conceptual Design (LVCD) method 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 ...
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The principle goal of this paper is developing of Launch Vehicle Conceptual Design (LVCD) method 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. Evaluation and verification of the presented method is performed using available data of two and three-stage launch vehicles.