Javad Vaziri Naein Nejad; F. Ommi; Seyed Hossein Moosavi
Volume 10, Issue 1 , June 2017, , Pages 35-45
Abstract
The addition of proper-sized metal particles to the effervescent fuels increases the density of exhaust gases from rocket engines and the trust consequently. On the other hand, the addition of non-optimized metal particles causes combustion instability. Thus, the separation of proper-sized metal particles ...
Read More
The addition of proper-sized metal particles to the effervescent fuels increases the density of exhaust gases from rocket engines and the trust consequently. On the other hand, the addition of non-optimized metal particles causes combustion instability. Thus, the separation of proper-sized metal particles is under consideration here. In this study, among different methods of separating the aluminum particles in the fuel, the performance of the conic cyclone separator has been studied and the numerical results are validated by the experimental data. With a specific particle diameter and speed, the less the angle between the cyclone body and the horizon, the higher would be the separation efficiency. In addition, for increasing the separation efficiency of aluminum particle, it is recommended to build the inlet section of cyclone at the lower point of cyclone body.
Space systems design (spacecraft, satellites, space stations and their equipment)
Hanieh Eshaghnia; Mehran Nosratollahi; Amirhossain Adami
Volume 14, Issue 4 , December 2021, , Pages 35-49
Abstract
A new approach to the design and development of launchers is the use of advanced technologies to reduce design and development costs as much as possible. In this paper, an approach to reduce costs and increase reliability is proposed, which is based on the use of a non-turbo pump propulsion system (pressure-fed ...
Read More
A new approach to the design and development of launchers is the use of advanced technologies to reduce design and development costs as much as possible. In this paper, an approach to reduce costs and increase reliability is proposed, which is based on the use of a non-turbo pump propulsion system (pressure-fed propulsion system) instead of a turbo pump propulsion system. For this purpose, the multidisciplinary conceptual design optimization of a two-stage launch vehicle with a pressure-fed propulsion system with the aim of sending max payload with a least gross mass to the orbit (500 km) in terms of structure, aerodynamics, propulsion, pressure vessels, simulation, and pitch program disciplines. Then, the sensitivity analysis was performed on the optimum launcher to determine the efficiency of the launcher at different orbital heights and the ability to carry a suitable payload.
Space systems design (spacecraft, satellites, space stations and their equipment)
Mahmoud Talafi Noghani; Peiman Aliparast
Volume 14, Issue 1 , March 2021, , Pages 35-41
Abstract
In this paper, a shielding box is designed for a Ku-band VSAT transmitter power amplifier based on an electromagnetic band gap structure. The proposed structure, is a bed of nails on the inner surface of the top wall of the box. It suppresses cavity resonances in the Ku-band uplink frequency range of ...
Read More
In this paper, a shielding box is designed for a Ku-band VSAT transmitter power amplifier based on an electromagnetic band gap structure. The proposed structure, is a bed of nails on the inner surface of the top wall of the box. It suppresses cavity resonances in the Ku-band uplink frequency range of 14-14.5 GHz. The amplifier which is composed of two cascaded modules (a pre-amplifier and a 50 W power amplifier) is first imported to a full wave simulation software. Full wave simulation of the printed circuit board (PCB) which includes the small signal scattering parameters of the power amplifier modules, proves the optimum performance of the proposed design.
Space systems design (spacecraft, satellites, space stations and their equipment)
Sajjad Davari; Hadiseh Karimaei; Mohammad Reza Salimi; Hassan Naseh
Volume 16, Issue 1 , March 2023, , Pages 35-46
Abstract
Monopropellant thruster are used to inject a satellite into orbit or control its position on three axes in space missions. One of them is hydrazine thruster which is widely used. In this research, design of the injector, decomposition chamber and nozzle of a 10N hydrazine monopropellant thruster have ...
Read More
Monopropellant thruster are used to inject a satellite into orbit or control its position on three axes in space missions. One of them is hydrazine thruster which is widely used. In this research, design of the injector, decomposition chamber and nozzle of a 10N hydrazine monopropellant thruster have been performed. The capillary injector was designed using Fluent software for this thruster which was able to supply the mass flow rate of the thruster (5 gr/sec). The decomposition chamber contains catalyst granules and its dimensions were selected based on the complete decomposition of hydrazine. The nozzle was designed by RPA software. The validation of the design with RPA software was checked by a numeric code. This code was able to calculate the dimensions of the decomposition chamber based on the amount of hydrazine decomposition. Accordingly, the results of both design methods are strongly consistent with each other. At the end of the design, the final thruster design and drawings were prepared.
Space systems design (spacecraft, satellites, space stations and their equipment)
Mostafa Jafari; Alireza Toloei
Volume 15, English Special Issue , May 2022, , Pages 35-44
Abstract
A numerical dynamic-aerodynamic interface for simulating the separation dynamics of constrained strap-on boosters jettisoned in the atmosphere is presented. Two commercial solvers: a 6DOF multi-body dynamic solver and a numerical time-dependent flow solver are integrated together with an interface code ...
Read More
A numerical dynamic-aerodynamic interface for simulating the separation dynamics of constrained strap-on boosters jettisoned in the atmosphere is presented. Two commercial solvers: a 6DOF multi-body dynamic solver and a numerical time-dependent flow solver are integrated together with an interface code to constitute a package that presents real-time dynamic/aerodynamic coupled analysis. Dynamic unstructured mesh approach is employed using local remeshing methods in respect of bodies motion with a second-order upwind accurate 3D Euler solver. This interface can simulate multi body separation dynamics interaction with aerodynamic effects to complete separation mechanisms like springs, thrusters, joints and so on. The flow solver is validated by the Titan IV launch vehicle experimental data. The separation integration is used for a typical launch vehicle with two strap-on boosters using spring ejector mechanism and spherical constraint joints acting in the dense atmosphere. Hence, the aim of the presented interface is to facilitate the integration of complicated separation mechanisms with a full numerical CFD aerodynamic solver.
M M. Moghaddam; A Salimi
Volume 1, Issue 1 , September 2008, , Pages 37-45
Abstract
This Paper presents a dynamic model of a micro-satellite in Mesbah class. At this model aerodynamic torque and solar radiation pressure torques are considered as disturbance torques. Gravity gradient torque is assumed as stabilizing torque and acts as a passive controller. Magnetic torquers act as an ...
Read More
This Paper presents a dynamic model of a micro-satellite in Mesbah class. At this model aerodynamic torque and solar radiation pressure torques are considered as disturbance torques. Gravity gradient torque is assumed as stabilizing torque and acts as a passive controller. Magnetic torquers act as an active controller. There are three methods of optimization of power consumption; first using LQR controller, secondly using the mapping function (which is suggested to ensure that the generated magnetic moment by the coils is perpendicular to the local magnetic field vector), and finally powering on control system over the earth stations only for the purpose of power saving.
E. Mohammadzaman; H. R. Momeni
Volume 3, Issue 1 , July 2010, , Pages 37-44
Abstract
In this paper a new guidance law is proposed to guarantee the stability of the guidance loop considering first order pursuit dynamics using short time stability theorem. As homing guidance is operates over a finite time, short time stability criterion which is defined over a specified time interval can ...
Read More
In this paper a new guidance law is proposed to guarantee the stability of the guidance loop considering first order pursuit dynamics using short time stability theorem. As homing guidance is operates over a finite time, short time stability criterion which is defined over a specified time interval can be used effectively in guidance loop stability analysis. Proposed guidance law utilizes line of sight angular rate and pursuit
acceleration measurements. Stability region which depends on the pursuit dynamics and guidance gains is an analytical expression in terms of time to go. Stability condition of the new guidance law is less conservatism than classical proportional navigation guidance law.
Ahmad Reza Sadeghi; Mohammad Farzan Sabahi; Sayed Mohamad Saberali
Volume 9, Issue 1 , May 2016, , Pages 37-46
Abstract
Automatic control of satellites and spacecrafts, has been extensively paid attention. Attitude determination is one of the most important procedures to control a spacecraft. Star trackers are widely used for attitude determining. A star tracker provides images from the around space and try to identify ...
Read More
Automatic control of satellites and spacecrafts, has been extensively paid attention. Attitude determination is one of the most important procedures to control a spacecraft. Star trackers are widely used for attitude determining. A star tracker provides images from the around space and try to identify the stars in the images. Several algorithms are proposed to this end. However, most of these algorithms use the raw measurements to star identification and attitude determination. As the measurements are often affected by various types of noise, the performance of such algorithms is degraded. Here, we employ tracking algorithms to improve the performance of existing methods for attitude determining. The Kalman filter-based tracking algorithms are shown to have satisfactory results for object tracking. We use the JPDAF to build an algorithm for accurate tracking of stars locations in successive images and, consequently, determining the attitude of spacecraft. The presented algorithm is compared with the well known algorithm for attitude determining called SNA.
Ashkan Mahmoud-Aghdami; Farhad Farhang Laleh; Mohammad Ghahramani
Volume 9, Issue 3 , December 2016, , Pages 37-51
Abstract
In this article the design and the development of spool type pin puller mechanism from concept model to engineering model was expressed. At first, present models and designs were investigated and the conceptual design was based on positive aspects of these models and in order to better study the operation ...
Read More
In this article the design and the development of spool type pin puller mechanism from concept model to engineering model was expressed. At first, present models and designs were investigated and the conceptual design was based on positive aspects of these models and in order to better study the operation of mechanism, the concept model was built 5 times bigger than the real model. After this model the pre-engineering model with 350N pulling force was designed and built. It was tried to reduce the operational defects of concept model like fuse operation and the rupture of warping wire, in the pre-engineering model. Assembling of the model was carried out with special tools and operation test carried out in room temperature. In the engineering model pulling force was increased to 500 N and the material of inside parts was changed due to special electrical and mechanical conditions. Temperature and atmospheric test conditions were considered in the last model and the model was prepared to pass all qualification tests. Vibration tests and also Temperature tests were performed in vacuum chamber with respect to space standards. Reliability calculations were done on each pin puller parts and at last reliability of whole pin puller system were obtained.
Space systems design (spacecraft, satellites, space stations and their equipment)
Hamidreza Soleimani; Milad َAzimi
Volume 13, Issue 4 , December 2020, , Pages 37-48
Abstract
This paper analyses the dynamic behavior of the rigid solar panels deploying mechanism of a spacecraft with flexible hinges. The proposed mechanism, maintaining a proper speed, guarantees the deployment synchronization of solar panels and minimizes the effects of impact and vibration applied during the ...
Read More
This paper analyses the dynamic behavior of the rigid solar panels deploying mechanism of a spacecraft with flexible hinges. The proposed mechanism, maintaining a proper speed, guarantees the deployment synchronization of solar panels and minimizes the effects of impact and vibration applied during the final stage and after the panels lock-up using torsional springs in the hinges and yoke driven assembly. The equations of the motion of the system are derived using Lagrangian approach and the behavior of the mechanism for constant and variable torque excitation modes is investigated. The simulation results presented along with the dynamic simulations performed by Adams software and conventional mechanisms show the efficiency of the proposed method.
communications
Peyman Mohammadi; Mehdi Alemi Rostami
Volume 17, Issue 1 , March 2024, , Pages 37-48
Abstract
One of the most important issues related to the power supply of TWTA lamps in satellites is to have low ripple, high efficiency, high reliability and optimal volume and weight,. In this article, the efficiency and reliability of high voltage DC/DC electronic-power converter is optimized for use in satellite ...
Read More
One of the most important issues related to the power supply of TWTA lamps in satellites is to have low ripple, high efficiency, high reliability and optimal volume and weight,. In this article, the efficiency and reliability of high voltage DC/DC electronic-power converter is optimized for use in satellite and TWTA lamps. The goal of optimization using multi-objective genetic algorithm (NSGA-II) in this article is to minimize the objective function, which includes efficiency and reliability. Markov model is used to evaluate reliability, in which short-circuit and open-circuit errors are considered for circuit switches and diodes, and short-circuit errors are considered for passive circuit elements. for optimization, first, the input variables of the algorithm are determined as the input of the objective function, so that with the help of sensitivity analysis, the parameters that have low sensitivity and their changes do not have a major impact on the objective function are eliminated. parameters of NSGA-II algorithm, including the number of iterations, the number of populations, and the probability have been determined for the accurate calculation of circuit variables. the results section this method, in addition to maintaining high efficiency, with the optimal selection of elements, high reliability can be achieved .
O. Shekoofa; M. Taherbane
Volume 2, Issue 2 , July 2009, , Pages 39-49
Abstract
This paper intends to study the impacts of orbit parameters change and evaluate their importance in Electrical Power Subsystem (EPS) design. Two main objectives have been followed in this research: 1) understanding the impacts of the orbital parameters change and the mechanisms of their interactions ...
Read More
This paper intends to study the impacts of orbit parameters change and evaluate their importance in Electrical Power Subsystem (EPS) design. Two main objectives have been followed in this research: 1) understanding the impacts of the orbital parameters change and the mechanisms of their interactions with the EPS design and operation, 2) evaluation of the importance of their effects. To this end, a typical cube satellite has been considered in different LEO orbits, to investigate the impacts of variation in the main orbit parameters e. g. altitude and inclination angle. Then the sizing, operation and performance of power sources have been evaluated via comparing the results of in-orbit simulations of EPS operation. In addition, some indirect impacts of the orbit parameters change are evaluated, by analysis and calculation of the interaction between EPS and other subsystems such as Telecommunication and Telemetery (TMTC), Attitude Determination and Control Subsystem (ADCS) and Thermal Control. The results show how the sizing and operation of solar array and battery are under the influence of orbit parameters change via certain factors such as orbit period, duration and the fraction of eclipse to sunlit phases, received solar irradiance by solar panels, and received thermal fluxes from the sun. According to the acquired results, any altitude increment leads to have better margins in power source sizing but there is an optimum value for inclination angle from this point of view.
F. Ommi; S. Askari Mahdavi; S. M. Hosein- Alipour; E. Movahed-Nezhad
Volume 4, Issue 1 , July 2011, , Pages 39-48
Abstract
A linear instability analysis of an annular liquid sheet emanating from an atomizer subjected to inner and outer air streams to investigate the liquid viscosity and swirl velocity on the maximum growth rate has been carried out. The dimensionless dispersion equation that governs the instability of a ...
Read More
A linear instability analysis of an annular liquid sheet emanating from an atomizer subjected to inner and outer air streams to investigate the liquid viscosity and swirl velocity on the maximum growth rate has been carried out. The dimensionless dispersion equation that governs the instability of a viscous annular liquid sheet under air streams was derived with linear stability analysis. The dispersion equation solved by numerical method and investigated viscosity and swirl velocity effect on maximum growth rate and its corresponding unstable wave number. The results show that decrease in viscosity has positive effect on maximum growth rate and its corresponding unstable wave number. At low liquid swirl Weber number liquid swirl has a stabilizing effect and at high liquid swirl Weber number liquid swirl velocity has a destabilizing effect on the liquid sheet. The growth rate can be related to the breakup length of the liquid sheet and when the growth rate increase, breakup length was shorter. The drop diameter dependent to the wave number and decrease with increase on it that afford to improvement the combustion and decrease the specific fuel consumption.
R. Kamali –Moghadam; S. Nouri; M. R. Salimi; M. Sheida; S. A. Hosseini
Volume 6, Issue 3 , October 2013, , Pages 39-48
Abstract
When a solver is used for analyzing the hypersonic reentry vehicles, high speed and accuracy of the solver results are the basic parameters in the design process. In the present study, the results obtained by solution of laminar boundary layer equations using integral matrix method and approximate method ...
Read More
When a solver is used for analyzing the hypersonic reentry vehicles, high speed and accuracy of the solver results are the basic parameters in the design process. In the present study, the results obtained by solution of laminar boundary layer equations using integral matrix method and approximate method are assessed in aeroheating prediction around hypersonic axisymmetric reentry bodies. The results show that the applied methods have suitable accuracy in aeroheating and high computational speed for reentry vehicle design. Space marching method in numerical simulation of boundary layer equations and applying less grid point in the boundary layer due to use of integral matrix method rather than other methods efficiently decrease computational costs. Also, high robustness of approximate method in the heat flux prediction over the reentry surface makes it useful for design process.Using a special approximate relation for stagnation region improves the aero-thermodynamics results.
M. Navabi; M. Tavana; H.R. Mirzaei
Volume 7, Issue 4 , January 2015, , Pages 39-49
Abstract
Attitude control of spacecraft in order to nonlinear and high order dynamics is fundamental and challenging issue. With respect to these nonlinearities, linear control theories are not suitable choices and spacecraft may be unstable or lose performance. In this paper, State Dependent Riccati Equation ...
Read More
Attitude control of spacecraft in order to nonlinear and high order dynamics is fundamental and challenging issue. With respect to these nonlinearities, linear control theories are not suitable choices and spacecraft may be unstable or lose performance. In this paper, State Dependent Riccati Equation (SDRE) method is utilized to 3-axis stabilization using four reaction wheels. State dependent Riccati equation method is systematic approach for optimal control of nonlinear systems which satisfies constraints of systems. In order to solve time consuming problem of this method in practical systems, Theta-D method is used. Results demonstrate the effectiveness of Theta-D in compare with Riccati method.
Reza Yavari; Iman Mohammadzaman; Mohammad Reza Arvan
Volume 9, Issue 4 , April 2017, , Pages 39-50
Abstract
In this paper, a novel integrated guidance and control (IGC) approach is designed using the combination of backstepping and sliding mode control methods. In contrast to the traditional methods combining the kinematic and dynamic equations and deriving a state space model as an integrated unit model, ...
Read More
In this paper, a novel integrated guidance and control (IGC) approach is designed using the combination of backstepping and sliding mode control methods. In contrast to the traditional methods combining the kinematic and dynamic equations and deriving a state space model as an integrated unit model, the proposed method designs the guidance and control problem in a single loop. This algorithm is robust with respect to the uncertainties in the target acceleration and missile dynamic model. Simulation results using six-degrees-of-freedom simulation aerodynamic model (6DOF) and three-dimension (3-D) engagement show that the proposed IGC design, with guidance and control dynamic synergism, eventuates interception with the maneuvering target.
Mohammad Amin Eskandari; Hasan Karimi; Davood Ramesh; Mohammda Reza Alikhani
Volume 13, Issue 1 , March 2020, , Pages 39-48
Abstract
Expansion cycle rocket engines have unintelligible and sensitive dynamic behavior. Contrary to other types of rocket engine which have gas generator, Expansion cycle rocket engines utilizes mass flow of fuel propellant to provide power for rotating turbo pump. Which contributes to a complicated and difficult ...
Read More
Expansion cycle rocket engines have unintelligible and sensitive dynamic behavior. Contrary to other types of rocket engine which have gas generator, Expansion cycle rocket engines utilizes mass flow of fuel propellant to provide power for rotating turbo pump. Which contributes to a complicated and difficult ignitions process in these engines. Priority and delay process in opening of control valves is important to prevent aforementioned phenomena. As opening and closing of control valves cause dynamic process in rocket engine, whose effects are expensive and difficult to predict by experimental tests. Therefore, dynamic modelling plays a key role in development of expansion cycle rocket engines and may decrees future expenses. In this article RL-10 rocket engine with sufficient data for validation has been chosen. The main goal of this article is dynamic modelling of expansion cycle rocket engine using mathematical non-linear models. Modelling results yield that the presented non-linear model is valid.
GPS and navigation GPS)، GLONASS، GALILEO
Mir Reza Ghaffari Razin; Behzad Voosoghi
Volume 13, Issue 3 , September 2020, , Pages 39-50
Abstract
In this paper, WNN with PSO training algorithm is used to modeling and prediction of time-dependent ionosphere total electron content (TEC) variations. 2 different combinations of input observations are evaluated. The number of stations used to train of WNN with PSO algorithm selected 20 and 10. In all ...
Read More
In this paper, WNN with PSO training algorithm is used to modeling and prediction of time-dependent ionosphere total electron content (TEC) variations. 2 different combinations of input observations are evaluated. The number of stations used to train of WNN with PSO algorithm selected 20 and 10. In all testing mode, 3 GPS stations with proper distribution are considered as a testing stations. Statistical indicators relative error, dVTEC and correlation coefficient were used to assess the wavelet neural network model. The results of proposed model compared with GPS-TEC and international reference ionosphere 2012 (IRI-2012) TEC. Average relative error computed in 3 test stations are 5.43% with 20 training station and 9.05% with 10 training station. Also the correlation coefficient calculated in 3 test stations are 0.954 with 20 training station and 0.907 with 10 training station. The results of this study show that the WNN with PSO algorithm is a reliable model to predict the temporal variations in the ionosphere.
Space systems design (spacecraft, satellites, space stations and their equipment)
Hojat Taei; Mahmood Haghighat Esfahani; Sajjad Yadegari Dehkordi
Volume 14, Issue 3 , September 2021, , Pages 39-50
Abstract
In this paper, a novel Comprehensive Preference-based Design (CPD) approach is presented which attempts to achieve subjective attributes that are defined in the concept of maximization of designer/customer's satisfaction in addition to objective goals which are formulated in the form of minimization ...
Read More
In this paper, a novel Comprehensive Preference-based Design (CPD) approach is presented which attempts to achieve subjective attributes that are defined in the concept of maximization of designer/customer's satisfaction in addition to objective goals which are formulated in the form of minimization of a performance criterion in a two-phase structure using two nested optimizers.In the first phase of CPD,using the concept of satisfaction,the subjective preferences of the designer/customer are defined in terms of fuzzy relationships and operators.Whereas the results of this phase are inaccurate,in the second phase,it is attempted to define a performance criterion and in order to achieve an optimal operational plan,attitude parameters and the compromises needed to meet the designer/customer's preferences are implemented.The methodology is utilized to design of a space launch vehicle for delivering 1200 kg payload to a 750 km orbit.Comparison of the results shows that despite the higher mass of launch vehicle designed by CPD,overall design satisfaction is higher and designer/customer's preferences have been satisfied.
J. Eskandari Jam; M. Noorabadi; S. H. Taghavian; N. Garshasbi-Nia
Volume 5, Issue 1 , April 2012, , Pages 41-50
Abstract
In this paper the mechanical behavior of satellite carrier adapter made of composite lattice shell is examined. First, the geometrical parameters of the composite lattice shell are analyzed. Choosing the direction for winding the fibers (geodesic route), geometric equations of the structure is elicited. ...
Read More
In this paper the mechanical behavior of satellite carrier adapter made of composite lattice shell is examined. First, the geometrical parameters of the composite lattice shell are analyzed. Choosing the direction for winding the fibers (geodesic route), geometric equations of the structure is elicited. Then, stiffness matrix of the structure is obtained according to these equations. Finally using finite element modeling of a conical lattice shell sample, the comparison between finite element and analytical results are presented. The analytical and numerical results show that with increasing rib’s thickness and Width, axial strain of the structure decreases nonlinearly.
Sohayla Abdolahi; Fskhri Etemadi; Mohammad Ebrahimi
Volume 8, Issue 3 , October 2015, , Pages 41-53
Abstract
In this study, the aerodynamic heating of the flying body during powered flight phase has been numerically investigated. The conjugate simulation of fluid heat transfer and solid heat conduction has been considered. To this aim, the coupling boundary condition has been used for body shell that allows ...
Read More
In this study, the aerodynamic heating of the flying body during powered flight phase has been numerically investigated. The conjugate simulation of fluid heat transfer and solid heat conduction has been considered. To this aim, the coupling boundary condition has been used for body shell that allows the conjugate heat transfer investigation in the fluid and solid domains simultaneously. The model has been considered as a circular cylinder and spherically blunted cone nose with 350mm in diameter. The investigation has been carried out at different Mach number from 1.5 to 4.2 to cover range of supersonic flow. The advantage of this method is that the wall temperature and heat flux ââin any part of the nose and body shell with or without axial symmetry, connected components and other protuberances could be calculated at different angles of attack. Finally, the approach has been validated through the results of analytical and numerical methods for aerodynamic heating of axisymmetric vehicles.
Farid Taji Hervi; ALIREZA NOVINZADEH
Volume 10, Issue 3 , December 2017, , Pages 41-57
Abstract
The purpose of the present paper is to prove the model-free optimal control theory. This theory is derived from the principles of dynamic programming and it is produced for discrete-time systems. The design of the controller depends merely on the I/O data of the controlled planet; hence, the controller ...
Read More
The purpose of the present paper is to prove the model-free optimal control theory. This theory is derived from the principles of dynamic programming and it is produced for discrete-time systems. The design of the controller depends merely on the I/O data of the controlled planet; hence, the controller is independent of the model. In this paper, two actions have been performed in order to measure the value of the controller. In the first step, the control method was designed to control the attitude of spacecraft. The purpose of this theory was to create a model-free optimal control for the spatial model and to measure the efficiency of the spacecraft systems. Secondly, designing linear quadratic regulator (LQR) controller for attitude control of spacecraft was carried out. The reason for designing this controller was to compare it with model-free optimal control. If the differences between two controllers was proved to be small, then the theory would be proven. Finally, it has been concluded that controller is valuable and acceptable.
mehdi alemi rostami; Morteza Aghaei
Volume 11, Issue 1 , June 2018, , Pages 41-48
Abstract
In this paper, a high efficiency DC-DC converter for variable input voltage and high output voltage applications is presented. This converter is specially appropriate for driving travelling wave tube amplifier with a variable input source (i.e. solar panels). The proposed converter consists of a boost ...
Read More
In this paper, a high efficiency DC-DC converter for variable input voltage and high output voltage applications is presented. This converter is specially appropriate for driving travelling wave tube amplifier with a variable input source (i.e. solar panels). The proposed converter consists of a boost converter to remove the variations of its input voltage and regulate its output voltage. Afterwards, a full bridge inverter controls the phase angle between the output inverter voltage and the current. Keeping the value of this angle at zero, the switching losses are reduced. A series-parallel resonant circuit uses the parasitic elements of the transformer as its elements and makes switches work in soft switching conditions. This reduces the converter power losses and increases the efficiency. Simulation results show the behavior of the proposed converter.
Amirhossain Adami; Hojat Taei; Mansour Hozuri
Volume 12, Issue 1 , April 2019, , Pages 41-53
Abstract
Considering the importance of the presence of uncertainties in the design of complex engineering systems, in this research multidisciplinary design optimization process for a bipropellant propulsion system in the presence of uncertainties, which in addition to minimizing the system mass, has a high robust. ...
Read More
Considering the importance of the presence of uncertainties in the design of complex engineering systems, in this research multidisciplinary design optimization process for a bipropellant propulsion system in the presence of uncertainties, which in addition to minimizing the system mass, has a high robust. Based on this, the multidisciplinary design view of the bipropellant propulsion system is expressed in both optimum design and optimum robust design. The continued with the application of uncertainties, the mass, operational and geometric results of the propulsion system are expressed in terms of optimum design, robust design and optimum robust design. According to the results, it is shown that the lowest mass occurs in optimum design mode. But with uncertainties, it is observed at this point that it has the least robust and reliability. It also attempts to explain the difference between the concepts of robust design and optimum design with the help of results
Space subsystems design: (navigation, control, structure and…)
Mohammad Haji Jafari; Afshin ValiMohammad; Mahsa Mahdavi
Volume 16, Issue 4 , December 2023, , Pages 41-56
Abstract
In this research, the dynamic response of a U-12 CubeSat (20x20x30 cm) is investigated for three configurations during the launch conditions. Despite the existence of a successful design for this satellite, adding a mount to install on a standard adapter, the structural design of the satellite has been ...
Read More
In this research, the dynamic response of a U-12 CubeSat (20x20x30 cm) is investigated for three configurations during the launch conditions. Despite the existence of a successful design for this satellite, adding a mount to install on a standard adapter, the structural design of the satellite has been modified for modal computability with the launcher., which is based on 3 general approaches: to apply geometric changes to the structure, changing the extent of the points to improve the limitations of connecting to the launcher, and rearranging the internal system components regarding the standards and principles of compatibility and proximity. Therefore, more than the addition of the adapter mount and a few changes in the separating plates, a spire has been added to the structure, which has caused a change in the internal arrangement, including the halving of the fuel tank (by reducing the capacity of 595 cubic centimeters). Despite the increase of 370 grams of total mass, the natural frequencies of the system have been increased enough without the need for redesign and there will be no frequency interference with the frequency spectrum of the launcher.
