نشریه علوم و فناوری فضایی
پیاپی 3 (بهار 1388)

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.

In this paper, a designing procedure of Single Gimbal Control Moment Gyro (SGCMG) for performing an agile slew maneuver in a microsatellite is described, then a prototype is fabricated and finally the test results are presented. The design of actuator mechanism is based on simplicity, direction of produced torque, minimum volume and weight. A DC electrical and a stepper motor with accuracy of 0.024 degree are used for controlling the angular velocity of flywheel and the gimbal slew rate, respectively. The motors controller and driver units are designed and implemented, so that the maximum accuracy, minimum errors and best response time could be accessible. The flywheel design is based on the required angular momentum which should be stored. The gimbal consists of two in-line beams which are attached to bearing in one side and momentum wheel system in the other side. A specific approach was considered to avoid any deformation in beams in consequence of mounting the momentum wheel system.

The paper presents a kinematical model for a random positioning system which simulates microgravity conditions on the ground. The random positioning system contains two frames, with an experiment platform. The frames are rotated around two perpendicular axes by a random angular velocity that it causes continuous changes in the orientation of experimental sample relative to the gravity’s vector. The angular velocity between random values is changed based on specific function in order to limit the generated spurious acceleration below a preselected threshold. The kinematical model yields criteria to evaluate the microgravity conditions on ground base. Such criteria are defined based on the mean values of acceleration and gravity components which are sensed by the rotating experimental sample.

In this paper, reliability growth concepts and modeling methods have been studied and a reliability growth model for launch vehicles has been proposed. Numerical example shows how we can obtain the model parameters. A method for calculating the initial launch vehicle reliability is suggested, also a method for estimating and validating of launch vehicle reliability in various number of launches with number of successes based on binomial distribution has been proposed. Results show after 10 successful launches we can conclude that launch vehicle reliability is between 0.741 and 1.

In this paper, equation of motion of three axis attitude dynamic of flexible spacecraft is derived using combination of finite element method and Euler equation. Flexible appendafes are modeled by beam elements. Goal of control is target attitude of spacecraft from initial state to desired attitude and suppression of vibration that induced in flexible appendages. So a combination of backstepping and sliding mode control method used for three-axis attitude maneuver of flexible spacecraft and for suppressing vibration of flexible appendage used from active vibration control method by PZT actuator. Control law for vibration control is based on LQG method.

Optimized injectors spray and consequently their proper fluid distribution is one of the most important factors determining the thrust force and combustion stability. Reaching a stable combustion with a high specific impulse needs a proper fuel distribution and atomization. Asymmetric distribution of fuel spray droplets lead to asymmetric spray cone and incomplete combustion that cause the chamber broken. Regarding to the numerous advantages of swirl double-based injectors, 25 number of this injector type are manufactured to be mounted on a circular injector plate. To ensure that they have enough accuracy, macroscopic spray characteristic of all injectors are measured and examined in cold test lab and PDA lab.Having checked the quality of each injector, two types of injector plates are designed and manufactured precisely. Having manufactured a circular and a hexagonally arranged injector plate, their spray characteristics are examined.Results show that the circular injector plate has satisfactory spray characteristics like droplet distribution and desirable spray cone and could be ready for warm test.

In this article the simulation of explosion and diffusion of debris consequent on explosion, in the space are studied. This method is simulated by the introduced code and the results are closely checked in accordance with different operational situations for the huge amounts of particles in different kinds of strategic orbits. Finally, the density distribution of debris that located in the same orbit is studied. The simulation examines the status of diffusion, trace settlement and also their location in orbits around the earth, during times that is wanted by user. Whatever satellite designing is immaculate, the satellite subsystems (such as batteries, propeller parts and etc) explosion and failing is possible, also smashing with the space masses is foreseeable. Regarding to the space orbits high expense and dedicating these orbits to each country; it is important to clean them from debris and necessary to understand this problem.

This article generates the algorithm for increasing performance of a space launch vehicle using solid rocket boosters.In all of the world, many space launch vehicles are designed with limited performance capabilities. But when the increased performance capabilities is required, one of the best solution that used on all of them, is using of solid rocket boosters. Solid rocket boosters has simple design and manufacturing capabilities and even converting the existing solid rocket booster blocks is one of the solutions.