A. Anvari; M. Shahriyari; F. Farhani
Volume 1, Issue 2 , December 2008, , Pages 1-8
Abstract
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 ...
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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.
SH. Marzban; K. Mohamedpour
Volume 2, Issue 1 , April 2009, , Pages 1-12
Abstract
Aeronautical telemetry system is applied to real flight conditions and movements so as to test the efficiency of different parts of an air vehicle such as an airplane, a missile, and a space shuttle during the flight. In order to study, determine and lessen the deleterious effects of the air vehicle's ...
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Aeronautical telemetry system is applied to real flight conditions and movements so as to test the efficiency of different parts of an air vehicle such as an airplane, a missile, and a space shuttle during the flight. In order to study, determine and lessen the deleterious effects of the air vehicle's flight maneuvers upon the aeronautical telemetry radio link, an urgent need for measuring the instantaneous communication angles between the telemetry receiver antenna and the transmitter antenna mounted on the air vehicle is highlighted. In the context of the values of these instantaneous communication angles, the gain from the telemetry receiver antenna and from the transmitter antenna throughout the flight trajectory can be obtained. It should be noted that in most previous studies, however, the gain from the receiver antenna during the test has been assumed to remain constant because the telemetry receivers have been provided with a system of auto-tracking the air vehicles. In order to study the improvements suggested by using coding, modulation, and other communication techniques and methods in the aeronautical telemetry radio link, a suitable model of real telemetry canal should be developed. By use of particular aeronautical navigation equations, the present paper is going to first develop an algorithm for measuring the instantaneous communication angles between the telemetry receiver antenna and the air vehicle during the flight. The equations are thus solved. After working out the equations, the value of the instantaneous power received by the receiver at any instant of the flight can be determined. On the basis of the suggested algorithm and through the simulation of the radio link (along the entire flight trajectory of an assumed air vehicle) the probability of the bit error rate of the collected data for some propagation environment of aeronautical telemetry can be found.
S. H. Pourtakdoust; R. Moradi; R. Kamyar
Volume 2, Issue 2 , July 2009, , Pages 1-16
Abstract
In this work the coupled nonlinear problem of optimal spacecraft rendezvous and docking (RVD) is addressed. In most of the previous studies on the subject of optimal RVD, decoupling is presumed to exist between the trajectory translational and the attitude motions and hence the optimal coupled analysis ...
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In this work the coupled nonlinear problem of optimal spacecraft rendezvous and docking (RVD) is addressed. In most of the previous studies on the subject of optimal RVD, decoupling is presumed to exist between the trajectory translational and the attitude motions and hence the optimal coupled analysis has not been yet addressed properly. However there are circumstances where these two motions are in fact coupled and interdependent and one such situation is investigated and analyzed in this article. By utilizing thrusters for the translational control and reaction wheels for the attitude control, one can uncouple the translational and rotational control to a high degree of approximation. However it can be shown that due to even very small thrust misalignments, the uncoupled problem changes to a highly coupled one. In this article, the nonlinear rendezvous and docking problem is assumed to be coupled and its optimal fuel-trajectory closed loop solution is obtained using two approaches of local linearization and Gauss Pseudospectral methods. Therefore the designed controllers are able to handle the highly nonlinear coupled rendezvous and docking optimally in the presence of system uncertainties as well as environmental disturbances. The results of the two solution approaches and their pertinent control strategies are compared and the merits and weaknesses of each are fully analyzed. Finally, a sensitivity analysis is also performed that shows the effects of thrust misalignments levels on the final state diversions.
S. H. Jalali-Naini
Volume 2, Issue 3 , December 2009, , Pages 1-12
Abstract
In this paper, a closed-loop optimal guidance with final position and velocity constraints is obtained by applying time-varying weighting coefficient in the performance index in order to shape the commanded acceleration. The control system is assumed to be linear, time-varying, and of arbitrary order ...
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In this paper, a closed-loop optimal guidance with final position and velocity constraints is obtained by applying time-varying weighting coefficient in the performance index in order to shape the commanded acceleration. The control system is assumed to be linear, time-varying, and of arbitrary order with a throttleable engine. The acceleration due to drag is also modeled as a linear function with respect to velocity vector multiplied by a given function of time. In addition, different weighting functions are suggested for different acceleration constraints, such as maximum dynamic pressure, separation of stages, and zero acceleration at the final time. Finally, the performance of the guidance law for a combined weighting function is evaluated and discussed.
M. Ahangar; R. Ebrahimi
Volume 3, Issue 1 , July 2010, , Pages 1-13
Abstract
In this study, the combustion process of gaseous Oxygen on the surface of HTPB solid fuel has been investigated. To simulate the chemically reactive flow, Navier-Stokes equations and species transport equations were solved using LU-SW implicit scheme. Modeling this kind of combustion process demands ...
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In this study, the combustion process of gaseous Oxygen on the surface of HTPB solid fuel has been investigated. To simulate the chemically reactive flow, Navier-Stokes equations and species transport equations were solved using LU-SW implicit scheme. Modeling this kind of combustion process demands a deep understanding of the pyrolysis phenomenon on the solid fuel surface. Experimental studies conducted in this field show that the main gaseous product of the pyrolysis process is C4H6. An experimental equation which is dependent to the temperature of the fuel surface is used to determine the gas production rate during pyrolysis process. The temperature of the fuel surface can be obtained by applying energy equation in gas-solid interface. The combustion process of gaseous Oxygen and C4H6 has been described by two quasi-global chemical kinetics models. According to the obtained results, the main characteristic parameters of combustive flow such as the flame temperature and mass fraction of chemical species are strongly affiliated to the applied chemical kinetic model. Finally, the results of modeling based on two different models of chemical kinetics are presented and solid fuel surface regression rate is compared with other numerical results.
Sh. Marzban; K. Mohamed-Pour
Volume 3, Issue 2 , January 2011, , Pages 1-9
Abstract
In the most aeronautical telemetry systems, at least two antennas are used to transmit radio signals towards receiver antenna. It is due to effect of large metallic fuselage of air vehicles in cutoff radio link between transmitter and receiver antenna during flight manoeuvres. Installation of two antennas ...
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In the most aeronautical telemetry systems, at least two antennas are used to transmit radio signals towards receiver antenna. It is due to effect of large metallic fuselage of air vehicles in cutoff radio link between transmitter and receiver antenna during flight manoeuvres. Installation of two antennas on the fuselage of air vehicle guarantees a convenient and continuous link between telemetry transmitter and receiver antennas. But during some moments that receiver antenna receiver radio signals from two transmitter antenna simultaneously, there is phenomena named self-interference, one can overcome this problem through making independence between two transmitting signals. In this paper using one of the block codes named Alamouti one can assure independence of two transmitting signals from two transmitter antennas. Using this code in a radio link, variation of bit error rate probability in flight path reduced to some convenient values and quality of radio link will be increased.
M. R. Ghaffari; M. M. Hossainali
Volume 4, Issue 1 , July 2011, , Pages 1-14
Abstract
The function based approach to three-dimensional tomographic modeling of Ionosphere is analyzed. Harmonic and empirical orthogonal functions are used as the base functions required in the modeling of the horizontal and vertical variations of the electron density. The instability of solution has been ...
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The function based approach to three-dimensional tomographic modeling of Ionosphere is analyzed. Harmonic and empirical orthogonal functions are used as the base functions required in the modeling of the horizontal and vertical variations of the electron density. The instability of solution has been numerically analyzed and the Tikhonov regularization technique has been used to regularize the solution. To come up with an optimum value for the regularization parameter, the direct measurements of the electron density obtained from the Tehran Ionosonde station, located at λ=51.640 and φ=35.870 , are used. The electron density model reconstructed in this paper has a maximum relative error of 36.44% and its minimum value is 0.8503. The maximum difference between the vertical total electron content (VTEC) obtained from the reconstructed model to that obtained from the corresponding IGS network Ionosphere product is 52.320 TECU and its minimum value is 1.268.
A. H. Adami; M. Nosratollahi
Volume 4, Issue 2 , January 2012, , Pages 1-10
Abstract
New algorithm is presented in this paper for attitude determination of LEO nanosatellite with 2 accuracy in attitude determination independent of time. The most important limitation in nanosatellites is about subsystems’s masses so, reduction of subsystems’s masses is always considered. ...
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New algorithm is presented in this paper for attitude determination of LEO nanosatellite with 2 accuracy in attitude determination independent of time. The most important limitation in nanosatellites is about subsystems’s masses so, reduction of subsystems’s masses is always considered. ADS plays the important role in the successful orbital maneuver missions. ADS accuracy is connected with increasing of sensors and complex processors which lead to increase the ADS mass. The presented algorithm uses one magnetometer sensor and one horizon sensor and position data receiving by GPS sensor as minimum required sensors. The selected configuration is resulted to minimum ADS mass and mission cost. Finally, error analysis at two most important orbit zones is done and the performance of the presented algorithm is confirmed.
F. Samadzadegan; Gh. Abdi
Volume 5, Issue 1 , April 2012, , Pages 1-14
Abstract
The increase in capability and performance of digital cameras, processors and image processing algorithms has caused vision-aided navigation of aerial vehicles to be a hot research of interest. In order to determine pose parameters form vision-aided navigation methods, it is common to use automatic image ...
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The increase in capability and performance of digital cameras, processors and image processing algorithms has caused vision-aided navigation of aerial vehicles to be a hot research of interest. In order to determine pose parameters form vision-aided navigation methods, it is common to use automatic image registration using information of reference databases. However, solving registration issue in automatic navigating of aerial vehicles has been considered a complex manner. In this paper, a novel method for vision-aided navigation of aerial vehicles to increase reliability and accuracy of geo-referencing aerial image is proposed. To have robust evaluation, different aerial images with variety of conditions are utilized to assess this method. Obtained results show high performance of proposed method to solve issues related to automatic GEO-referencing of aerial images.
S. S. Nasrolahi; H. Bolandi; M. Abedi
Volume 5, Issue 2 , July 2012, , Pages 1-13
Abstract
In this study, a fault tolerant Attitude Determination System (ADS) has been designed which provides fault detection, isolation and tolerant abilities in this system. Suggested approach is based on derivation of all possible rotations between body and orbital frames and comparison of Euler angles provided ...
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In this study, a fault tolerant Attitude Determination System (ADS) has been designed which provides fault detection, isolation and tolerant abilities in this system. Suggested approach is based on derivation of all possible rotations between body and orbital frames and comparison of Euler angles provided by them. In this regard, significant changes in the variance of Euler angles set are considered as criteria for fault detection. Moreover, fault isolation and tolerant mechanisms are based on classification of rotation matrices which are not affected by faulty components. The above features present a quite analytical and computational approach which does not impose additional mass, power consumption and cost in the satellite. Also, designed diagnosis and fault correction algorithms are model-free basedmechanisms which always provide tolerated attitude angles for the attitude control subsystem. The mentioned abilities combined with the model based FDI mechanisms utilized in the attitude control system, provide an advanced decision support system capable of isolation of faults which have been simultaneously occurred in the satellite sensors and actuators. Finally, performance of the designed algorithm is approved by simulation results.
Reza Omidi Gosheblagh; Karim Mohammadi
Volume 5, Issue 3 , October 2012, , Pages 1-9
Abstract
Due to high design and launch cost of satellites, their failure probability should be minimized. Single Event Effects (SEUs) are one of the most common error sources in satellite microelectronic. To cope with these unwanted errors, various techniques are used. The reliability analysis of these methods ...
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Due to high design and launch cost of satellites, their failure probability should be minimized. Single Event Effects (SEUs) are one of the most common error sources in satellite microelectronic. To cope with these unwanted errors, various techniques are used. The reliability analysis of these methods is one of the major acceptance criteria to validate these techniques. In order to evaluate the reliability of satellite subsystems, it is required to determine the SEU rate as a primary factor. A practical method to determine this rate is based on Weibull approach in which the SEU cross section is used as an initialized parameter. In this paper, the SEU rate is calculated based on weibull method for Low Earth Orbit (LEO) satellites, as case study Iranian demonstrated Rasad and Omid satellites. Furthermore, based on the proton density, an accurate time-varying SEU rate model is proposed which determines the rejuvenation time for SEU susceptible subsystems.
Jafar Roshanian; Mehdi Hassani; Shabnam Yazdan; Masoud Ebrahimi
Volume 5, Issue 4 , January 2013, , Pages 1-8
Abstract
Star tracker is an attitude determination device which determines the satellite or spacecraft’s attitude using the star position information in inertial and body references. Star information is collected and stored onboard as a “Star or mission catalog”. There are several star catalogs ...
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Star tracker is an attitude determination device which determines the satellite or spacecraft’s attitude using the star position information in inertial and body references. Star information is collected and stored onboard as a “Star or mission catalog”. There are several star catalogs that contain different kinds of information with different accuracy. In this paper the most used star catalogs are introduced and a few star catalog selection features are recommended. These features are weighted according to the star tracker mission type. For the selected star tracker mission, results demonstrate that Hipparcos star catalog is the best choice. Eventually using Hipparcos star catalog, a mission catalog is developed to be used onboard a typical star tracker.
F. Moosavi; J. Roshanian; R. Emami
Volume 6, Issue 1 , April 2013, , Pages 1-10
Abstract
This paper presents the control design for large angle and high rotation rates maneuvers using reaction cold gas thrusters. Navigation system provides suborbital attitude changes in terms of quaternion. Cold gas thrusters with pulse-width pulse-frequency modulation provide nearly proportional control ...
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This paper presents the control design for large angle and high rotation rates maneuvers using reaction cold gas thrusters. Navigation system provides suborbital attitude changes in terms of quaternion. Cold gas thrusters with pulse-width pulse-frequency modulation provide nearly proportional control torques. The use of quaternion as attitude errors for large angle feedback control in a suborbital capsule is investigated. Numerical simulations demonstrate the practical feasibility of a three-axis large angle maneuver.
A. A. Mehmandoost-Khajeh-Dad; M. Khaghani; M. Jafarzadeh-Khatibani
Volume 6, Issue 2 , July 2013, , Pages 1-10
Abstract
Being a very powerful method, Positron Annihilation Spectroscopy (PAS) has been widely used for investigation of defects type and concentration in materials in recent years. In this paper, we first report characteristics of a Positron Annihilation Lifetime Spectroscopy (PALS) system which has been made ...
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Being a very powerful method, Positron Annihilation Spectroscopy (PAS) has been widely used for investigation of defects type and concentration in materials in recent years. In this paper, we first report characteristics of a Positron Annihilation Lifetime Spectroscopy (PALS) system which has been made in an Iranian nuclear spectroscopy instrument company with corporation of Sistan and Baluchestan University for the first time. The system is working with a resolution better than 350ps and can be used for materials such as ceramics, glasses and insulator materials. Next, we report designing and development of the first slow positron beam in Iran and we explain its advantages for exploring defects in materials. Positron Source, moderator and vacuum chamber including positron beam tube and sample chamber, has been designed and prepared. Designing of extraction lenses and magnetic fields has been made using CST STUDIO software in order to optimum focusing and transition of beam.
H. R. Ali Mohammadi; D. Ramesh; M. R. Heidary; R. Farrokhi; H. Karimi
Volume 6, Issue 3 , October 2013, , Pages 1-13
Abstract
In this paper, a particular propulsion system including, liquid rocket engine, fuel and oxidizer tank and related pressurizing system, have been surveyed. The procedure is based on a nonlinear mathematical model which has been simulated in Matlab Simulation environment. In propulsion systems, identifying ...
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In this paper, a particular propulsion system including, liquid rocket engine, fuel and oxidizer tank and related pressurizing system, have been surveyed. The procedure is based on a nonlinear mathematical model which has been simulated in Matlab Simulation environment. In propulsion systems, identifying system performance is essential, because if we can accept ability describe the dynamic behavior of the system components in nominal and transient regimes, we can reduce the associated costs during design and development. Following, results of Propulsion system hot test are compared with model that shows acceptable accuracy of simulator code. In addition to leading research, how to use this model to identify the causes of failure is shown. Match analysis and compatibility testing, after disassembling objective observations show considerable performance model for similar applications.
M. Shafiey Dehaj; R. Ebrahimi; H. Karimi; A. Jalali; M. Naderi
Volume 6, Issue 4 , January 2014, , Pages 1-11
Abstract
The effects of liquid propellant rocket engines thrust termination transients are important in achieving the launch vehicle’s desired final velocity with the required precision. In this paper, a mathematical model has been developed to predict the changes in the combustion chamber pressure and ...
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The effects of liquid propellant rocket engines thrust termination transients are important in achieving the launch vehicle’s desired final velocity with the required precision. In this paper, a mathematical model has been developed to predict the changes in the combustion chamber pressure and the related cut-off impulse based on the physical aspects of engine’s components. The modeling is divided into four steps: (1) from issuing the cut-off command to the activation of valve (2) from the time of executing the stop command until the end of the operation of the cut-off valves, (3)from the time that after the valve is shutt closed until the combustion chamber is extinguished and (4) simulation of the phase change and propellant components evaporation in corrugates. Results suggest that the duration of the first two steps have a significant effect on increasing or decreasing the amount of the thrust force and the 4th step’s thrust is less than 10 percent of nominal value while, the most thrust fluctuations appear in this step.
A. Jafarsalehi; M. Mirshams; R. Emami
Volume 7, Issue 1 , April 2014, , Pages 1-12
Abstract
This paper focuses upon the development of an efficient method for conceptual design optimization of a satellite. There are many option for a satellite subsystems that could be choice, as acceptable solution to implement of a space system mission. Every option should be assessment based on the different ...
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This paper focuses upon the development of an efficient method for conceptual design optimization of a satellite. There are many option for a satellite subsystems that could be choice, as acceptable solution to implement of a space system mission. Every option should be assessment based on the different criteria such as cost, mass, reliability and technology contraint (complexity). In this research, mass and technology constraints, which have a direct impact on the satellite life cycle cost, are considerd as system level objective function to obtain the system optimal solution during the coceptual design phase. The approach adopted in this paper is based on a distributed collaborative optimization (CO) framework. At system level, multiobjective optimization goal is to minimize the dry mass of the satellite and, simultaneously, minimize the system technology complexity which is subject to equality constraints. The use of equality constraints at the system level in CO to represent the disciplinary feasible regions, introduces numerical and computational difficulties as the discipline level optima are non-smooth and noisy functions of the system level optimization parameters.To address these difficulties robust optimization algorithms such as genetic algorithms (GA) are used at the system level. The results show that the CO framework has the same level of accuracy as the conventional All-At-Once approaches.
A.A Nikkhah; J. Tayebi; J. Roshanian
Volume 7, Issue 2 , July 2014, , Pages 1-9
Abstract
In this paper attitude control system of nanosatellite based on Single Gimbal Control Moment Gyroscope (SGCMG) is presented. A LQR/LQG method is developed for stability of satellite and a feedback quaternion strategy is used for maneuvering mode. In the stabilization mode LQR/LQG controllers ...
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In this paper attitude control system of nanosatellite based on Single Gimbal Control Moment Gyroscope (SGCMG) is presented. A LQR/LQG method is developed for stability of satellite and a feedback quaternion strategy is used for maneuvering mode. In the stabilization mode LQR/LQG controllers are designed with linearization of nonlinear dynamic equation of satellite and control moment gyroscope, so that in other reseach didn’t use this controller in the stabilization mode of this system. In the maneuvering mode a feedback quaternion controller applyed for nonlinear system. Numerical simulations are provided to show the efficiency of the proposed controller for a nanosatellite with four single gimbal control moment gyroscope pyramid cluster. Results of simulations shown that LQR/LQG method is more accurate in compared with feedback quaternion controller.
S. M. SalehiAmiri; A. A. Nikkhah; H. Nobahari
Volume 7, Issue 3 , October 2014, , Pages 1-8
Abstract
This paper presents a method for calculation the non observable states in alignment and calibration process in gimballed inertial navigation system, using estimation method in static linear system and heuristic optimization algorithms. The non observable constant states in alignment process are horizontal ...
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This paper presents a method for calculation the non observable states in alignment and calibration process in gimballed inertial navigation system, using estimation method in static linear system and heuristic optimization algorithms. The non observable constant states in alignment process are horizontal accelerometers biases and azimuth gyroscope drift. In order to use the estimation method in static system, the observations are recorded in necessary time duration to convert the dynamic alignment process to static process. Simulation results show appropriate accuracy of purposed method for calculation the non observable states. Although the case study is the alignment process for gimballed inertial navigation system, the purposed method can be used for calibration and alignment of any inertial navigation systems.In purposed method the genetic heuristic optimization algorithm is used.
Seyed Hamid Jalali Naini
Volume 7, Issue 4 , January 2015, , Pages 1-9
Abstract
In this paper, implicit guidance equations are derived in polar coordinates. Depending on applications, implicit guidance equations in polar coordinates may be preferred over cartesian coordinates. Moreover, depending on the type of guidance problem, analytical solutions for sensitivity matrices may ...
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In this paper, implicit guidance equations are derived in polar coordinates. Depending on applications, implicit guidance equations in polar coordinates may be preferred over cartesian coordinates. Moreover, depending on the type of guidance problem, analytical solutions for sensitivity matrices may be simplified using polar coordinates. Therefore, transformation of implicit guidance equation into polar coordinates can be useful in guidance problems. In addition, the resulting equations are extended to cylindrical coordinates.
Tahere Binazadeh; Mohammad Hossein Shafiei; Elham Bazregarzadeh
Volume 8, Issue 1 , April 2015, , Pages 1-7
Abstract
This paper presents a novel approach in design of missile guidance law against highly maneuvering targets. This approach is based on the principles of partial stability and finite-time stability (finite-time partial stability). Also, it is shown that the designed guidance law is in conformity with a ...
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This paper presents a novel approach in design of missile guidance law against highly maneuvering targets. This approach is based on the principles of partial stability and finite-time stability (finite-time partial stability). Also, it is shown that the designed guidance law is in conformity with a real guidance scenario that leads to collision. In the design procedure the acceleration vector of the target is assumed as an external bounded disturbance and only this bound is required in the design of the guidance law. Therefore, the maneuver of the target is not restricted to any known and predetermined structure and measurement or estimation of the target acceleration vector during the maneuver is not necessary. The performance of the proposed guidance law is shown through analysis and computer simulations.
Ahmad Izadipour; Bhzad Akbari; Ali Reza Sharifi; Meysam Yousefzade
Volume 8, Issue 2 , July 2015, , Pages 1-10
Abstract
Comparison of direct and indirect methods for georeferencing of satellite linear array stereo images is proposed in this paper. In direct method, collinear condition and orbital elements are used for geolocation of pixels and ground control points (GCPs) are used in indirect method. After geolocation ...
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Comparison of direct and indirect methods for georeferencing of satellite linear array stereo images is proposed in this paper. In direct method, collinear condition and orbital elements are used for geolocation of pixels and ground control points (GCPs) are used in indirect method. After geolocation of initial points, a polynomial transformation is used to change pixels coordinates into corresponding geographical values. We validate these methods through experiment with SPOT-4 and IRS-P6 images. Experimental results indicate that indirect method can be used for georeferencing of satellite linear array images and the accuracy is as much as or even better than that of the direct method but it is time- and cost-consuming and also needs operator.
Ali Reza Alikhani; Yosef Shamadi
Volume 8, Issue 3 , October 2015, , Pages 1-13
Abstract
Important issues in designing a controller for re-entry vehicles is environmental uncertainties such as rapid changes in atmospheric properties which is an explicit function of altitude and also uncertainties of itselfvehicle such as aerodynamic coefficient, moment of inertia and so on. This paper deals ...
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Important issues in designing a controller for re-entry vehicles is environmental uncertainties such as rapid changes in atmospheric properties which is an explicit function of altitude and also uncertainties of itselfvehicle such as aerodynamic coefficient, moment of inertia and so on. This paper deals with the design of a control in order to overcome the uncertainty thatuses bank angle as a trajectory control variable.Another issue raised in recent studies has been integration of adaptive controller with guidance systems of re-entry vehicles because in real re-entry vehicle the bank angle is not a predefined profile function of velocity or altitude buta guidance algorithm are usedto produce bank commands during the atmospheric flight. Hence, other objectives of the thesis is to study and implementing of a guidance algorithm and proving of desired performance of the designed controller in a perfect scenariofrom starting point of the re-entry path until the opening of parachutes. Performance of designed controller is studied through simulations ofsix degrees of freedom of re-entry vehicle.. The results showed good performance in the presence of parametric uncertainty andunknown initial condition.
Ali Reza Alikhani; Seyed Aliakbar Kasaeian
Volume 8, Issue 4 , January 2016, , Pages 1-7
Abstract
Tracking guidance commands for a time-varying aerospace launch vehicle during the atmospheric flight is considered in this paper. Hence, the dynamic terminal sliding mode control law is constructed for this purpose and dynamic sliding mode control is utilized. The terminal sliding manifold causes the ...
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Tracking guidance commands for a time-varying aerospace launch vehicle during the atmospheric flight is considered in this paper. Hence, the dynamic terminal sliding mode control law is constructed for this purpose and dynamic sliding mode control is utilized. The terminal sliding manifold causes the dynamic sliding mode to converge asymptotically to zero in finite-time. The actuator and rate gyro dynamics are included in the model of launch vehicle. Dynamic sliding mode control accommodates unmatched disturbances, while the terminal sliding mode control is used to accelerate the system to reach the dynamic sliding manifold. Finally, the effectiveness of the proposed control is demonstrated in the presence of unmatched disturbances and is compared with the dynamic sliding mode.
Davood Ramesh; Sajad Khodadadiyan; Hasan Karimi
Volume 9, Issue 1 , May 2016, , Pages 1-11
Abstract
The purpose of this paper is to present a genetic algorithm (as a software) to optimize engine main parameters through the application of "genetic algorithm" and also introduced the new and modified thermodynamic cycles with analysing their performance. This software objective function is to achieve ...
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The purpose of this paper is to present a genetic algorithm (as a software) to optimize engine main parameters through the application of "genetic algorithm" and also introduced the new and modified thermodynamic cycles with analysing their performance. This software objective function is to achieve the highest and optimum level of 'final velocity'. In this study, the strategy of using fuel booster turbopump and 2nd stage fuel pump is followed primarily to moderate the effect of cavitation on pumps. Although the use of boosterpumps increase the weight, arise pumps' rpm and possibility to reduce the tanks pressure came with a decrease in weight of propulsion system. The developed software is applied to Russian RD-180 engine in construction of propulsion system of first stage of ATLAS IIIB LV, and experimental results have been demonstrating the improvement of engine performance which results from a multi-variable sensitivity study on a staged-combustion engine will be highlighted. This algorithm is under the limitation of constraints to control the critical variation of combustion pressure, turbine rpm, and pumps cavitation margin and turbine temperature. Results show that, supply flow rate of gas generation from 2nd stage of fuel pump and divide flow rate of exhaust of fuel booster turbine to 2nd stage of fuel pump and combustion chamber, will increase the final velocity of launch vehicle.
