نشریه علوم و فناوری فضایی
سال پانزدهم شماره 3 (پیاپی 53، پاییز 1401)

To create drag and reduce the speed of space payloads in the phase of entering the atmosphere, the payload body itself can be used as brake mechanisms without using additional tools. The approach analyzed in this paper is the separation of the nose and then the stability of the cylindrical body in horizontal or vertical mode. First, bynumerical solution, the cylindrical body is aerodynamically simulated in the flight conditions entering theatmosphere, and the location of the center of mass is designed to achieve static stability. Then, by developing the equations of motion of atmospheric reentry using aerodynamic coefficients and derivatives calculated by DATCOM, the flight parameters for both modes are compared and evaluated. The simulation results show that the horizontal flight is more efficient and is able to create better conditions for opening the parachute and landing. Another advantage of atmospheric reentry flight in horizontal mode is the proper distribution of aerodynamic heating and reduction of heat load in certain points of the payload.

In this paper, in addition to investigation and analyzing the dynamic model of a maneuver target, a new method based on the Interaction Multiple Model (IMM) method is presented to solve the tracking problem in presence of measurement noise. In this procedure, two models are used along with an extended Kalman filter for each model, for estimation of the states related to stochastic target model. To this end, a specific weight is calculated adaptively for each model and the final estimation of the target is obtained from the weighted sum of the modes related to each model. In this paper, second order Markov models are used to better describe the system behavior which leads to a decrease in the number of required motion models. This means that the previous two models are used to decide on the next model, and a much better algorithm is provided than the first-order IMM algorithm.

Deep learning is a modern method of image processing and data analysis that has entered the field of urban management with promising results and high potential. The purpose of this study is to investigate data augmentation techniques in improving the results of segmentation of building using aerial images with high spatial resolution and deep learning method. For this purpose, MSB building data set and MapNet model were used. The model was trained and evaluated in three stages without data augmentation, with data augmentation of geometric transformations and with data augmentation of geometric and photometric transformations. The results of model evaluation showed that using geometric transformations as data enhancement techniques, F-1 and IoU score evaluation criteria have increased by 0.5 and 0.55%, respectively, and using data techniques Incremental geometric and photometric transformations increased by 1.41 and 1.57 percent. This increase was visually observed in the improvement of the segmentation of dense areas of the building and the discontinuity of large-scale buildings.

In this paper, by creating and developing a code based on thermodynamics and gas dynamics equations, the performance characteristics of a 1N hydrazine monopropellant thruster such as thrust force, specific impulse, characteristic exhaust velocity, and propellant mass flow rate have been studied theoretically in terms of reaction chamber temperature. In this regard, by taking into account the adiabatic assumption, the reaction chamber temperature of monopropellant thruster has been analyzed zero-dimensionally using the ammonia dissociation rate as an independent variable under equilibrium and non-equilibrium conditions and it has been analyzed one-dimensionally using the hydrazine and ammonia homogeneous and heterogeneous reaction rate constants. Also, the effect of nozzle throat thermal expansion on reaction chamber pressure, thrust force, and propellant mass flow rate and the effect of reaction chamber pressure on ammonia dissociation rate and consequently on reaction chamber adiabatic temperature under thermodynamic equilibrium conditions have been studied.

In the present work, the frequency-domain aeroelastic stability analysis of space launch vehicle body in the flight condition of initial launch phase is presented for a range of geometric parameters, structural characteristics, and other parameters such as thrust force. The aeroelastic model is derived using structural assumed modes and quasi-steady aerodynamics. The pressure distribution of subsonic flow on the 3D configuration is determined by boundary element method. Non-uniform Euler-Bernoulli beam including torsion spring junctions along the body with free-free ends is used to model the structure, and its modal analysis is performed by finite difference method. Concluded results illustrate variation in parameters not only could vary the aeroelastic instability boundary, but also might cause the instability type changed (from divergence to flutter), which its main reason is replacement the second instability of the aeroelastic system with the first one. Furthermore, it is demonstrated that the follower thrust force restricts the aeroelastic stability but maintains the instability type.

Proportional navigation is one of the most widely used methods in guiding flying objects. This method requires the rotation rate of the line between the interceptor and the target to calculate the guidance command. For a variety of reasons, including cost savings, simple sensors are used to measure tracking information, including line of sight angle. Therefore, some non-measurable information such as the angular velocity of the line of sight must be estimated using mathematical equations. Due to the noise and other problems, the use of derivatives is not desirable in this situation. Therefore, in this paper, an extended nonlinear observer is used to estimate the angular velocity of the line. Due to the nonlinear dynamics of the intercepting of flying objects, a nonlinear type of observer has been selected. By performing a computer simulation, the correct operation of the proposed observer is shown.

The purpose of this paper is to present the cost estimation and optimization of space propulsion systems. Thus, choosing optimal propulsion system (from fuel and oxidizer aspect) is done in order to increase the efficiency and decrease the cost. Also, human resource cost and technology development time based on the consideration of labor cost effect on the personals motivation have been optimized. To this end, cost estimation and optimization algorithm has been drawn and suggested. The suggested algorithm has two steps. The first step in the algorithm is concern to cost estimation for seven fuel and oxidizer components. In the second step, labor cost and project implementation time is estimated and optimized based on the optimal space propulsion system derived from the previous step. Here, the objective functions are propulsion system technology development cost and time. On the other hand, the purpose is to consider the salary enhancement and consequently efficiency enhancement, time decrease and cost decrease.

In this paper, the feasibility and performance of using solar arrays equipped with sun concentrators, along with other conventional solar array structures, in CubeSats, is investigated for the first time. For this purpose, seven different structures of solar arrays have been defined and implemented for more than 24 different CubeSat configurations from 0.25U to 27U. Then, by calculating important system-level parameters such as power generation density, power generation cost, reliability of solar arrays, and also a newly proposed parameter, called shape fit factor, the performance of these structures for the introduced configurations are evaluated and compared. To this end, and by considering rational coefficients, a cost function consisting of the four above-mentioned parameters is defined as the degree of merit of different solar array structures used in each CubeSat configuration. The results show that alongside the use of deployable solar arrays, using concentrating solar arrays can provide new capabilities for CubeSats to overcome the challenge of generating sufficient power.

In this research, design and simulation of a single capillary injector and three-hole circular injector plate of a 10N Hydrazine monopropellant thruster were performed. Ansys Fluent software was used to simulate the injector and injector plate . Volume of fluid (VOF) method was used to simulate such a flow and turbulence was simulated by k-e model. The characteristics of the injector and injector plate including mass flow rate and average velocity in the injector nozzle were calculated by changing the inlet pressure. The results showed that the injector and the injector plate have the ability to supply the desired mass flow rate of the monopropellant thruster at a known design pressure. In fact the capillary injector has replaced swirl injector with hollow cone spray used in the previous version of this thruster. The dimension of the chamber was significantly reduced by using the capillary injector, which reduces both the volume of the expensive iridium catalyst and weight of the thruster.