نشریه علوم و فناوری فضایی
سال سوم شماره 3 (پیاپی 7، پاییز و زمستان 1389)

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.

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 in the small satellite community and 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 that the issues relating to 􀈝PPT 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 􀈝F, 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 􀈝N-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.

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 10th generation 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.

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.

Online optimal reentry guidance of reentry vehicles is the main objective of this paper. The procedure is based on the Matched Asymptotic Expansion (MAE) method, one of the Singular Perturbation Theory (SPT) procedures, and is aided with the Variation of Extremals (VOE) method. The new technique, named MAEOG(Matched Asymptotic Expansion Optimal Guidance) offers a very low solution time and an acceptable accuracy compared with the other numerical methods used until now for reentry optimization. Furthermore, it permits considering both the lift and the aerodynamic roll angle as control variables. The features of the new method appear completely suitable to develop a guidance scheme for atmospheric reentry.

The coupled rigid-body/slosh/elasticity dynamics equations are developed for 6-DoF flight of launchers. The equations of motion are derived by means of Lagrange’s equations in terms of quasi-coordinates and alternatively in the inertial frame. The simple pendulum model for planar motion is extended to model slosh dynamics in 6-DoF flight and the elastic motion is represented in terms of modal displacement coordinates relative to the elastic mean axes system. It is shown that this model is consistent with the simpler model for planar motion which has been developed in previous studies. The proposed dynamics model is incorporated in conjunction with the models for the other subsystems in a MATLAB/Simulink program to simulate 6-DoF flight of launchers.