نشریه علوم و فناوری فضایی
سال ششم شماره 4 (پیاپی 17، زمستان 1392)

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.

Solar panels are used in satellites to absorb solar energy and supply the power needed for space missions. Mission definition and satellite lifetime is restricted by the way and the amount of energy that can be supplied for satellite subsystems. Therefore, the design of satellite solar panels as a unique source for power supply in satellites is very sensitive and important. In this paper, an efficient design for solar panels of a GEO communication satellite is proposed considering the maximum strength to weight ratio and minimum deflection. For this purpose different space structures including Iso Grid structures, honeycomb and composite plates are studied. Several structural analyses are performed on the models in order to assurance from the strength and durability of the model in space working environments. The design according to composite-honeycomb is introduced as the best model for solar panel structure. The proper configuration of layers is also presented by developing a code based on an optimization algorithm.

Shape and size of sensitive volume are the most important parameters to model electronic devices for calculation of the SEU rate from space radiations. So far different models have been proposed for estimation of the sensitive volume. In this work, results of three models including RPP, Tetrahedral and Nested in calculation of sensitive volume have been compared with experimental result for AT60142 SRAM. GEANT4 as a Monte Carlo code has been used to calculate energy loss and energy straggling of ions with considering metallization and oxide layers. Comparisons between Monte Carlo and experimental results shows that RPP model estimates the SEU cross section with a large deviation in whole LET range, tetrahedral has good response in low LET's but dont follow experimental result for high LET particles and nested sensitive volume produce acceptable results for whole of LET range.

An explicit guidance law is developed for a reentry vehicle. Motion is constrained to a three-dimensional Bezier curve. Acceleration commands are derived by solving an inverse problem that combined with differential flatness approach. Trajectory is related to Bezier parameters. A comparison with pure proportional navigation shows the same accuracy, but a higher capability for optimal trajectory to some degree. Other advantages such as trajectory representation with minimum parameters, applicability to any reentry vehicle configuration and any control scheme, and Time-to-Go independency make this guidance approach more favorable.

Stress environment of space flights including hypergravity and microgravity effect cell functions and processes. We studied effect of hypergravity and flight stress on function, death, growth and cellular damage of five cell lines including human vascular endothelial cells, rat bone marrow stromal cells, mouse embryonic fibroblast cells, PC12 cells and pancreas islands in Biological payload of 3th sounding rocket by NO, LDH, MTT and insulin assay kits. Our results indicated that effect of hypergravity increase significantly cellular death and NO secretion but we found no changes in insulin secretion and pancreas function.

According to mission requirements in many of space missions, an exact time should be assigned for launch date namely launch window. Among various parameters influencing launch window, in this article a couple of most mission related ones are being researched. First is some constraints on handling satellite power and the second is the imaging mission constraintsfor a remote sensing satellite. Then after this study on mission requirements a conceptual method for launch window estimation is presented and supported by a case study on sample micro-satellite IRSAT.

In this paper, a multi-objective guidance scenario has been developed which could be used to improve the accuracy of sub-orbital modules at the entry to the earth’s atmosphere. Developed algorithm could be used for the trajectory shaping and adjusting of the trajectory relative to the desired position. To achieve defined objectives, it has been organized in three phases relative to different layers of trajectory. In the first phase, using of sensitivity functions which are of less computational burden, improving the accuracy according to the reference trajectory is done. In the second phase, trajectory shaping which is required to achieve desired flight performance is applied which could be designed and performed online in the mission. Finally in the last phase which is relative to low altitudes, improving of the accuracy according to homing rules is performed. Validation of developed algorithm using of a sample guidance problem is presented finally.

Rapid growth in using Commercial of The Shelf (COTS) components in Space missions compelled the researchers to evaluate the performance of this components in space environment. AT90CAN128 commercial microcontroller is used in AutSat mission due to high performance, simple structure, low power consumption and the ability of CAN bus management. In this paper the result of Total Ionizing Dose test on this microcontroller is presented. These results confirm the possibility of using this microcontroller for our satellite with minimum 3 years life cycle.