نشریه علوم و فناوری فضایی
سال پانزدهم شماره 1 (پیاپی 51، بهار 1401)

In the present study, the effects of geometrical properties of gas flow channels on both current density and temperature distributions inside a polymer electrolyte membrane (PEM) fuel cell are investigated. The main purpose here is to clarify the effects of the variation of width, depth, and the ribs of flow channels on the fuel cell performance. To do this, the fuel cell is numerically simulated in two dimensions. The governing equations consist of the conservation of the electrical potential, Darcy’s law as alternative to the momentum equation, Maxwell-Stefan equation for mass transport, energy conservation, and electro-thermal equations along with the Butler–Volmer equation. Numerical results indicate that the width of channels and their ribs have more sensible effects than the depth of flow channels on the current density and temperature distributions and fuel cell performance. While the maximum temperature of the cell is increased by increasing the width of the flow channels, the current density distribution and fuel cell performance can be improved. By decreasing the width of their ribs or depth of channels, the performance of the fuel cell is improved and its maximum temperature is decreased.

In this research, three-dimensional and four-dimensional tomography is used to demonstrate the distribution of wet refractivity index of the troposphere. In this model, spherical cap harmonics are used for the horizontal distribution of the wet refractivity index, and empirical orthogonal functions are used for the vertical distribution of the index. The region of study is in the west California State, and the wet refractivity index is retrieved from the wet tropospheric delay measurements. to validate the results, radiosonde profiles were compared to the tomographically retrieved profiles. The result shows that wet refractivity indices can be retrieved using functional models with RMSE about 2.4 ppm till 3.9 in four-dimension method. The comparisons show that the four-dimensional retrieved profiles shows improvement up to 34 and 42 percentage in mid-day tomography epochs compare to three-dimensional tomography results. Also it can be seen that in mid-night epochs three-dimensional tomography has higher accuracy compare to four-dimension method because of low variation of wet refractivity indices

Superhydrophobic coatings can be made by creating a micro-sized structure on a surface providing super-repellent properties. Numerical simulation of drop dynamics and motion on a superhydrophobic surface helps us understand control and building surface textures and find optimum micro structured coatings of maximum hydrophobicity. In the present work, the dynamics of drops on superhydrophobic inclined micro-structured surfaces is studied, using a finite element method. The governing equations and important dimensionless numbers are described and a numerical algorithm is introduced. The validation of the numerical algorithm is performed by simulation of drop motion attached to an inclined surface. In addition, droplet movement on the micro structured surface is numerically simulated on smooth and microstructured surfaces in the same conditions. Comparison of the results shows the effect of microstructured coating on the surface hydrophobicity properties.

A numerical dynamic-aerodynamic interface for simulating the separation dynamics of constrained strap-on boosters jettisoned in the atmosphere is presented. Two commercial solvers: a 6DOF multi-body dynamic solver and a numerical time-dependent flow solver are integrated together with an interface code to constitute a package that presents real-time dynamic/aerodynamic coupled analysis. Dynamic unstructured mesh approach is employed using local remeshing methods in respect of bodies motion with a second-order upwind accurate 3D Euler solver. This interface can simulate multi body separation dynamics interaction with aerodynamic effects to complete separation mechanisms like springs, thrusters, joints and so on. The flow solver is validated by the Titan IV launch vehicle experimental data. The separation integration is used for a typical launch vehicle with two strap-on boosters using spring ejector mechanism and spherical constraint joints acting in the dense atmosphere. Hence, the aim of the presented interface is to facilitate the integration of complicated separation mechanisms with a full numerical CFD aerodynamic solver.

Diversity plays an essential role in increasing reliability of redundant systems, however in safety and mission critical applications. The onboard computer of satellites and the flight computer of aircrafts, which are ultra-reliable system, utilizes various hardware platforms for their redundant architecture to resolve the common cause failure problem. Furthermore, the software is also developed by separate teams based on different software platforms to mitigate the specification and design flaws, and implementation mistakes. This paper focuses on modelling diversity of redundant systems using Markov reliability analyzing method. The proposed scheme is explored in two types of applications: mission critical applications (with long mission time) and safety critical applications (with short mission time). Analytical and simulation results show the effectiveness of diversity in increasing the reliability of these systems. Since about ten percent of all failures appear as common cause failures, which restrict reliability improvement through similar redundant modules, achieving ultra-reliable necessitate to consider diversity in these systems

The cold gas thruster is one of the significant components of a satellite and its application possesses a marked impact on the entire system performance. The nonlinear function and order of magnitude, lead to increasing the importance of thruster function. Therefore, pre-mission performance assessment has a considerable effect on the risk reduction of space missions. In this article, an uncomplicated and efficient pendulum scheme for development and implementation of a Thruster Test Stand (TTS), to measure the thrust produced at the end of the nozzle is proposed. The TTS is capable of measuring thrust levels in the range of 0.1Newtons to 3N with operating frequencies up to 50 Hz which is used by various satellite ranges. The experimental results demonstrate that although the designed device is less sophisticated than other test devices, it is capable of measuring the produced thrust very precisely and with less than 15mN.