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. Nosratollahi; A. h. Adami-Dehkordi
Volume 3, Issue 2 , January 2011, Pages 11-22
Abstract
This paper presents the multidisciplinary design optimization of monopropellant propulsion system of the nanosatellite for planner maneuver. Mass, configuration and internal ballistic equations are derived for any part of propulsion system (thruster, tank, pressurized gas, ...). Minimizing total mass ...
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This paper presents the multidisciplinary design optimization of monopropellant propulsion system of the nanosatellite for planner maneuver. Mass, configuration and internal ballistic equations are derived for any part of propulsion system (thruster, tank, pressurized gas, ...). Minimizing total mass of the propulsion system and satisfying all constrains such as Thrust limitation 5 (N) and 10 (N), Minimum specific impulse () and minimum throttle area (). AAO framework is developed and the direct search is selected for optimization method. Finally optimum designs are introduced and compared for 10(N) and 5(N) monopropellant propulsion system.
A. Rezaiha; M. Anbarloi; M. Farshchi
Volume 3, Issue 2 , January 2011, Pages 23-30
Abstract
Although Pulsed Plasma Thruster (PPT) has first been utilized in a space mission in 1964 but after more than four decades, it is still a space rated technology which has performed various propulsion tasks from stationkeeping tasks to three-axis attitude control for a variety of former missions. With ...
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Although Pulsed Plasma Thruster (PPT) has first been utilized in a space mission in 1964 but after more than four decades, it is still a space rated technology which has performed various propulsion tasks from stationkeeping tasks to three-axis attitude control for a variety of former missions. With respect to the rapid growth inthe small satellite communityand the growing interest for smaller satellites in recent years, PPT is one of the promising electric propulsion devices for small satellites (e.g. CubeSats) as the following advantages: simplicity, lightweight, robustness, low power consumptions, low production costs and small dimensions. In spite of the fact thatthe issues relating to μPPT scaling have been investigated to a certain degree in recent years, it is felt that for an application on CubeSats this topic has to be investigated in greater detail for even smaller dimensions and better performance. Therefore a laboratory benchmark rectangular breech-fed pulsed plasma thruster (PPT) was designed, developed and successfully tested in a bell-type vacuum chamber at 10-6 mbar for the first time in west Asia (Iran). The PPT has been tested while the main capacitor, which is a 35 μF, 2.5 kV oil-filled capacitor, has been charged with a wide range of voltage, ranging from 250 V to 1750 V making the system stored energy range from less than 1 J to 60 J, producing the impulse bit varying from 30 μN-s to 1.3 mN-s. This work initiated a research program in Iran for working on PPTs and miniaturization of PPTs while increasing the performance parameters. The present paper reviews the PPT design and the development briefly.
M. Naderi Tabrizi; S. A. R. Jalali Chimeh; H. Karimi
Volume 3, Issue 2 , January 2011, Pages 31-43
Abstract
In this article the Propellant Utilization system (PU) has been simulated. The objective of this system is to adjust the consumption ratio of the propellants in order to ensure the minimum propellant residuals at engine’s shutdown phase. Using the PU system, the orbital or range and also the payload ...
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In this article the Propellant Utilization system (PU) has been simulated. The objective of this system is to adjust the consumption ratio of the propellants in order to ensure the minimum propellant residuals at engine’s shutdown phase. Using the PU system, the orbital or range and also the payload capabilities of missiles or launch vehicles will be enhanced. In this article, after studying and simulation of the PU system, the necessity of using such system is compared with a missile without the PU system. At the end of this paper it is proven that using PU system on a desired missile has enhanced its range up to 7 percent and has also reduced the propellant residuals up to 25 percent.
M. Navabi; N. Nasiri
Volume 3, Issue 2 , January 2011, Pages 45-52
Abstract
Since last decades utilizing satellites in low earth orbits have had increasing tendency. These satellites experience the earth magnetic field due to their low altitude to the earth. The Earth magnetic intensity can be used in order to control the attitude of spacecraft utilizing the interaction between ...
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Since last decades utilizing satellites in low earth orbits have had increasing tendency. These satellites experience the earth magnetic field due to their low altitude to the earth. The Earth magnetic intensity can be used in order to control the attitude of spacecraft utilizing the interaction between the earth magnetic field and magnetic dipoles which are generated in the body of satellite. First of all, for using this phenomenon the magnitude and direction of the Earth magnetic field have to be obtained. There are various ways in order to simulate the earth magnetic field, that the most accurate one is utilizing the harmonic coefficients and mathematical model of the earth magnetic field. In this study, the earth magnetic field is modeled based on the 10thgeneration of the IGRF coefficients and the results are verified with the most valid reference. Due the Earth magnetic field is used in order to attitude control of a spacecraft, it is necessary to transform the results into the spacecraft Body frame. This transformation can be obtained utilizing linear and nonlinear transformation. In the next step, based on the comparison of the results of the spacecraft attitude dynamics utilizing linear and nonlinear transformation the validity margin of linear transformation is studied.
M. Jafari Nadoushan; M. Tivay
Volume 3, Issue 2 , January 2011, Pages 53-58
Abstract
Effect of regularization on the solution of perturbed two body problem is investigated in this paper. Purposes of using this method are computational burden reduction and achieving desirable accuracy in the minimum time. In this regard the equations of motion are linearized and independent variable is ...
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Effect of regularization on the solution of perturbed two body problem is investigated in this paper. Purposes of using this method are computational burden reduction and achieving desirable accuracy in the minimum time. In this regard the equations of motion are linearized and independent variable is changed from time to the true anomaly. These yield reducing run time, however increasing accuracy. The results of simulation confirm that utilizing this method in onboard computation or long term simulations is more suitable and efficient than other methods including general and special perturbation methods.
E. Amani; M. Ebrahimi; J. Roshanian
Volume 3, Issue 2 , January 2011, Pages 59-67
Abstract
در این مقاله، معادلات کوپلشده جسم صلب- تلاطم سیال- الاستیسیته برای پرواز شش درجه آزادی ماهوارهبرها توسعه داده شده است. معادلات حرکت به کمک معادلات لاگرانژ در دستگاه ...
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در این مقاله، معادلات کوپلشده جسم صلب- تلاطم سیال- الاستیسیته برای پرواز شش درجه آزادی ماهوارهبرها توسعه داده شده است. معادلات حرکت به کمک معادلات لاگرانژ در دستگاه شبه مختصات و همچنین در دستگاه مختصات اینرسی استخراج شدهاند. مدل پاندول ساده برای حرکت صفحهای به منظور مدل کردن دینامیک تلاطم سیال در پرواز شش درجه آزادی گسترش داده شده و تغییر شکلهای الاستیک بر اساس مختصات مودال نسبت به مختصات میانی ارائه شدند و نیز نشان داده شده است که این مدل با مدل سادهتر حرکت صفحهای که در مطالعات پیشین توسعه یافته است، سازگاری دارد. مدل دینامیکی پیشنهاد شده در کنار مدلهای لازم برای سایر زیرسیستمها در برنامة متلب/ سیمولینک با موفقیت برای شبیهسازی حرکت شش درجه آزادی ماهوارهبرها به کار گرفته شده است.
M. Mortazavi; D. Abbasi-Moghadam
Volume 3, Issue 2 , January 2011, Pages 69-76
Abstract
هدف اصلی مقاله حاضر، هدایت بهینه و برخط اجسام بازگشتی به زمین است. روند دستیابی به این مهم مبتنی بر روش بسط مجانبی هماهنگ است که یکی از روشهای خانوادة اغتشاشات تکین است ...
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هدف اصلی مقاله حاضر، هدایت بهینه و برخط اجسام بازگشتی به زمین است. روند دستیابی به این مهم مبتنی بر روش بسط مجانبی هماهنگ است که یکی از روشهای خانوادة اغتشاشات تکین است و به کمک روش تغییر اکسترمالها تقویت شده است. روش جدید حاصل MAEOGکه مخفف کلمات مربوط به هدایت بهینة مبتنی بر بسط مجانبی هماهنگ است ضمن ارائة راه حل با دقت قابل قیاس با روشهای دیگر، بسیار سریع مسئله را به جواب میرساند و زمان حل را کاهش میدهد. به علاوه، این امکان را میدهد که چه برا و چه زاویة رول به عنوان متغیرهای کنترل در نظر گرفته شوند. ویژگیهای روش جدید برای توسعة الگوریتم هدایتی بازگشت به زمین کاملاً مناسب به نظر میرسند.