Journal of Space Science and Technology
Year:2020 | Volume:13 | Issue:1 (42)|Papers Count:9

Using linear programming method in control allocation for attitude control subsystem of spacecraft with redundant thrusters is studied in this paper. The simplex algorithm is utilized as a solver and the Direction Preserving and Bodson’ s Reduced size Direction preserving approaches are used as optimal approaches to deal with non admissible solutions. Also, proper functioning of these approaches against thrusters fault phenomenon is evaluated. The results show that the Direction Preserving approache has less computational time and less fuel consuming. However, the Bodson’ s Reduced size Direction preserving approache has more computational time and more fuel consuming but less tri-axis tracking error. It should be noted that the PD controller has been used as a spacecraft control rule, and simulations have been made for the number and configuration of the specific thrusters.

In this research, it is attempted to determine the diameter and velocity distribution according to the flow characteristics of the upstream and without needing experimental measurements. Firstly, Fluent software has been used to simulate the turbulent flow of inside nozzle by k-? model to obtain the nozzle turbulence energy at the nozzle outlet. Then, nonlinear growth rate analysis of instability is used to determine spray breakup length and the frequency of maximum instability and the mean diameter of primary breakup. Four equation maximum entropy model has been developed according to the inlet of upstream flow. Subsequently, the terms of momentum source as well as the energy of maximum entropy model has been determined using the results of simulated nozzle turbulence flow and instability analysis. In the following, first the results of maximum entropy model have been evaluated with the experimental input and then determined with upstream input. The obtained results which have been compared with experimental tests show well agreement.

In the present paper, the process of designing launch abort system is presented for manned spacecraft in sub-orbital mission. For this purpose, by studying the other manned spacecraft launch abort system, a number of dimensionless parameters have been generated for the launch abort system and spacecraft. using these parameters, the mass, geometric-dimensional characteristics, the launch escape motor thrust and acceleration of the LAS have been estimated. Also The numerical aerodynamic analysis has been used to form the external configuration, and the results of the analysis, along with other effective parameters, have created a design framework. AHP analysis has been used to select the optimal configuration and in order to validation, an analytical model based on the momentum equation used to design the launch escape motor. Comparison of the results of this model and the statistical design has shown that the data produced is correct.

By incorporating provisions in Space Law documents, gateway for commercial space activities was opened. From the era of space treaties, commercial uses always are getting faster Due to the increasing of players and development of new technologies. Although the current framework does not comply with the legal issues arising from rapid and ongoing process, the provisions of these instruments have so far been the basis for action by governments & Non-governmental Orgs. in carrying such activities. This paper assuming the need to further development of space law in terms of its commercial dimension, but intends to consider the possibility or refusal of commercial activities and the status of soft law through international cooperation and development of national space legislation to relative normality of space. An initial review indicates the prescription of these activities, despite the cessation of contractual normalization and existence of constraints derived from principles such as the prohibition of appropriation, the common heritage of mankind & equitable sharing.

Expansion cycle rocket engines have unintelligible and sensitive dynamic behavior. Contrary to other types of rocket engine which have gas generator, Expansion cycle rocket engines utilizes mass flow of fuel propellant to provide power for rotating turbo pump. Which contributes to a complicated and difficult ignitions process in these engines. Priority and delay process in opening of control valves is important to prevent aforementioned phenomena. As opening and closing of control valves cause dynamic process in rocket engine, whose effects are expensive and difficult to predict by experimental tests. Therefore, dynamic modelling plays a key role in development of expansion cycle rocket engines and may decrees future expenses. In this article RL-10 rocket engine with sufficient data for validation has been chosen. The main goal of this article is dynamic modelling of expansion cycle rocket engine using mathematical non-linear models. Modelling results yield that the presented non-linear model is valid.

In this paper, the performance of fluid momentum controller (FMC) actuators in satellite temperature management is investigated based on two pyramidal and 3-axis proposed configurations. In this regard, the temperature of different satellite surfaces with fluid actuators and without actuators in an orbital period of satellite is investigated and the results are compared to each other. For FMC actuators that are closed as a loop, a Moving Reference Frame (MRF) is used and the flow inside the actuators is laminar. The effect of fluid angular velocity of actuators on the temperature of satellite surfaces with two different angular velocity has been investigated and time-dependent heat flux is applied to the satellite surfaces. The results indicate that in the pyramidal configuration, the decrease in the temperature of the satellite wall surfaces is influenced by two parameters: fluid angular velocity and orbital period of satellite, but in the critical conditions, the 3-axis configuration can carry out this temperature management more quickly.

Amount of space exploration investigation has been done to understand the variations in biological structure and function of living organisms in microgravity condition. Nowadays, the investigation of the microgravity from a nanoscale viewpoint is encouraged. Silver nanoparticles have been involved in large-scale production. In this research, the effect of simulated microgravity on DNA structure was studied. Additionally, the silver nanoparticles binding with DNA molecules under Earth gravity and simulated microgravity conditions by various spectroscopic instruments were investigated. The results displayed that microgravity simulation has created DNA structure variation. The binding affinity of silver nanoparticles to DNA altered. Microgravity initiated an alternation in size and surface charge of DNA and modified DNA structure from B to C-form. Consequently, based on our observation, microgravity can strictly affect the silver nanoparticles-DNA binding interaction. Our observations can open fascinating research lines in biology and biophysics.

The need to stay in space and use its advantage in different perspectives, has led our country to take steps to use satellite technology. In this regard, constructing an advanced satellite launch site with all the modern facilities is one of the key issues. Imam Khomeini Space Center is the first permanent launch platform of the Islamic Republic of Iran. In 1389, it was announced that due to existing geographic constraints, research was conducted to build the second space center in Chabahar. In this paper, using several performed studies on the infrastructure of the advanced space centers around the world as well as various simulation methods, a conceptual design for the Chabahar launch site is presented. Full descriptions of the launch platforms, processing complex, assembling of satellites, etc. are illustrated. The design and construction of the Chabahar launch site will be a major step forward in the development of the space industry of Islamic Republic of Iran.