فهرست مطالب حسن ناصح

Achieving to new technologies with high reliability, along with reducing the cost and time of the design cycle, is one of the most important challenges in complex systems. In this paper, reliability based design of a space system is discussed in the conceptual design phase. Normally, there are eight steps in the design for reliability. The first step, planning, and the next seven steps that applied for the liquid propellant engine with electro-pump technology included: determination of failure modes; reliability modeling; reliability allocation; propagation of uncertainty; Implementation of the chosen method in reliability analysis; reliability prediction and reliability evaluation. Therefore, in this research has been performed to achieved method and implementation steps of reliability design in the conceptual design phase of a space system.

The purpose of this article is to optimize the power supply subsystem of the satellite based on mass and power production. For this purpose, the satellite orbit is to determine the time of placement in the shadow in Satellite Tool kit (STK) software, and the design of the satellite energy supply subsystem of existing analytical equations are done, and these main design parameters have been optimized to achieve the minimum weight of system. Then, the beginning of the target satellite, the Orsted satellite, the orbital and mission operational modes of this satellite has been compiled and the statistical information of its power supply subsystem has also been determined. Then, the design of the power supply sub-system is based on the parametric analysis and processing of statistical information and also based on the simulation and orbital analysis of the satellite. Finally, in order to model the design problem, the objective functions of design and optimization, the weight of the power supply subsystem has been considered, and the main design parameters affecting the objective function are capacity, depth of discharge, efficiency and density of the battery and the area of ​​the arrays. It is worth mentioning that the type of battery itself is also very effective on the objective function, but considering the type of satellite battery considered (nickel-cadmium), the variable battery type is not considered. Finally, this problem has been optimized using genetic algorithm. The obtained results show the accuracy of the method and implementation compared to the values ​​of the existing satellite power supply subsystem.

This research aims to model the nonlinear dynamics of the space propulsion system equipped with an electro-pump. For modeling, this system includes the combustion chamber, fuel and oxidizer tanks, and the corresponding pressurization system. A set of governing equations has been modeled in the Simulink of MATLAB software. The results of the modeling including changes in the angular velocity versus the time, changes in combustion chamber pressure versus the time, and changes in the mass flow rate of the fuel and oxidizer to the combustion chamber versus the time, are presented. In the end, the obtained results are compared with the results of the available engine and in addition to validating the results, it is possible to confirm and validate the performance of the electro-pump in a liquid fuel engine. The results showed that the angular velocity rate, the average acceleration rate of the combustion chamber, the average flow rate of the oxidizer to the combustion chamber and the average flow rate of the fuel to the combustion chamber obtained from the simulation and the available reference have a maximum error of 20%. The amount of difference in the conceptual design stage is free of problems. In addition, the trend of the graphs is similar.

In this paper, the catalyst bed of a 10 N hydrazine monopropellant thruster was designed. The catalyst bed is including iridium granules, which is used to decompose the hydrazine in monopropellant thruster. Hydrazine must be decomposed almost completely in the catalytic chamber, because it is a carcinogenic chemical fuel and on the other hand, achieving the maximum power from the thruster is also an important goal. As a result, the effect of change in catalytic chamber length on the mass fraction of chemical species including hydrazine, ammonia, nitrogen, and oxygen was studied. Also, after determining the length of the catalytic chamber, the diameter of the nozzle throat corresponding to the same length was determined.

Monopropellant thruster are used to inject a satellite into orbit or control its position on three axes in space missions. One of them is hydrazine thruster which is widely used. In this research, design of the injector, decomposition chamber and nozzle of a 10N hydrazine monopropellant thruster have been performed. The capillary injector was designed using Fluent software for this thruster which was able to supply the mass flow rate of the thruster (5 gr/sec). The decomposition chamber contains catalyst granules and its dimensions were selected based on the complete decomposition of hydrazine. The nozzle was designed by RPA software. The validation of the design with RPA software was checked by a numeric code. This code was able to calculate the dimensions of the decomposition chamber based on the amount of hydrazine decomposition. Accordingly, the results of both design methods are strongly consistent with each other. At the end of the design, the final thruster design and drawings were prepared.

The main purpose is to introduce the performance system design and optimization method of aerospike nozzle for different aero-space conditions. For this purpose, some of the important parameters of the aerospike nozzle structure and cold flow condition tests in the nozzle optimization are studied. The methods of designing the Aerospike nozzle and its governing equations are described and the proposed design model is described and important factors are expressed in this type of nozzle. therefore, the design of a complete nozzle is made by aerospike and is supported by an existing design sample. Then, in order to optimize the nozzle, three cuts of 20%, 40% and 60% of the nozzle end are analyzed. The standard for comparison and optimization in these three slices is the Mach number of the output current. The results of this comparison show that the most efficient aerospike nozzle is a 40% cut nozzle based on the flow charts and contours of this aerospace nozzle.

The purpose of this study is to provide a method to prevent unwanted vibrations in environmental tests of random vibrations of spacecraft. Unwanted vibrations in many environmental random vibration tests will lead to the hidden fault (hiding failure) that would be happened in operational conditions in a space mission. Hence, in this research, first, a method called the power spectrum density notching method was presented to prevent unwanted vibrations, and then, by using the obtained analytical relationships, environmental random vibration tests of a sub-orbital spacecraft were performed. After the test, the recorded results by the fixture and payload's accelerometer sensors have been analyzed. In this analysis, the performance of the power spectrum density notching method was examined. The obtained results indicate, although the power spectrum density notching method is a simple, practical, and operational method but can be an efficient method to prevent unwanted vibrations to spacecraft and also prevent damage to spacecraft during testing.

Accurate solving of complex systems such as space systems and specifically space propulsion system is very costly and time consuming. By developing and building a surrogate model, the solution time and the cost can be reduced. The closer the surrogate model is to the actual model, the more accurate the solution and the lower the error rate. High-precision successor models are called metamodels. The basis of producing a high-precision meta-model is to perform high-precision sensitivity analysis with a suitable method. Sensitivity analysis can show the effect of input variables on output variables and produce a surrogate model by eliminating ineffective input variables. Therefore, sensitivity analysis is of great value in solving complex systems. The purpose of this paper is to analyze the sensitivity of the multidisciplinary design of a monopropellant liquid propulsion system by the Latin Hypercube Sampling method. In this article, the topics related to the liquid monopropellant propulsion system are divided into six parts: High pressure gas tank, liquid fuel tank, injector, decomposition chamber, catalytic bed and nozzle. By determining the input and output variables of each subject, the results of sensitivity analysis are displayed in two ways: the sensitivity of the input variables to the output and the two-by-two correlation of the parameters with each other. In the results, as can be seen, the specific impulse input variable, in the high-pressure gas tank and the liquid fuel tank, has no effect on the output variables. In the injector, the number of grooves, groove angles and fuel tank pressure do not have a significant effect on the output variables. In the decomposition chamber sensitivity analysis diagram, the radius of the granule and for the catalyst bed, in addition to the radius of the granule, the percentage of ammonia decomposition are also ineffective. Finally, the sensitivity analysis for the nozzle shows that the ratio of specific heat has no effect on the output variables.

In this paper, the structural optimization of a space capsule has been discussed by approximating a thin-walled cylindrical shell with a certain length under the axial compression force and constant lateral pressure. Design variables include the outer diameter and cylinder thickness. The purpose of optimization is to minimize the mass and maximize the frequency of the first vibration shape mode of the cylinder. Design constraints include the buckling load multiplier (buckling safety factor) above 1.5 and Von Mises stress below 100 MPa. In this problem, first, according to the permissible limits of the design variables, a design of experiment and then a sensitivity analysis have been carried out to check the sensitivity of the objective functions and constraints to the design variables. After numerically solving the output values with the help of Ansys software and preparing the response surface, the optimal design point has been identified with the help of the Genetic algorithm. Then, with the numerical simulation of the optimal point, the accuracy of the values obtained from the response surface method was checked and their accuracy was confirmed. It has also been observed that at the selected design point, Von Mises stress is less than its allowed value, i.e. 100 MPa, and also the buckling load factor is more than twice its minimum allowed value. However, this point has the smallest distance from the origin and the optimum point has been chosen as the knee point.

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.

When designing complex products such as space thrusters, accurate simulation models are needed to evaluate and improve the design during development. The implementation of these accurate simulation models is often expensive and time-consuming. Surrogate models or metamodels are simplified models of accurate and expensive simulations that can be used to reduce some computational costs during studies or design optimization. The closer the surrogate model is to the real model, the more accurate the solution and the lower the percentage of error. These models with high accuracy are called metamodels. The purpose of this article is to design the metamodel of the liquid single-propellant thruster system using the kriging method, which can predict the behavior of the model to some extent. The purpose of this article is metamodeling of liquid monopropellant propulsion system by kriging method, which can predict the behavior of the model to some extent. Metamodel and distribution diagram of design points related to each of the subjects are extracted. In addition to the mass metamodel of each of the discipline, for the injector, the metamodel related to the mass flow rate of the fuel, for the catalytic bed, Also, four Gaussian, Exponential, Linear and Spherical functions with degree two were compared for each of the meta-models in the kriging method. In this comparison, it was observed that due to the same coefficient of determination, the Gaussian function has less error than other functions and, as a result, better accuracy.

One of the most important steps in designing spacecraft is designing their control system. There are many types of control systems, often including electric thrusters, cold gas thrusters, and monopropellant thrusters. Monopropellant thrusters generate more propulsion than other thrusters and are more cost-effective for short space missions. In this study, arrangement of 12 hydrazine monopropellant thrusters on a spacecraft was investigated. Based on the required torque, twelve thrusters were used to control the situation of the spacecraft, so that each axis is controlled by four thrusters and each thruster must provide the propulsion force equal to 10 N. For this purpose, 10 N hydrazine monopropellant thrusters were used. Finally, three different arrangements of thrusters were presented. The third type of arrangement was selected due to its independence from changes in the space center of mass of the spacecraft and also zero slope of the thrusters relative to the main axes.

The purpose of this paper is to present the cost estimation model for Cryogenic/Semi-Crogenic space propulsion systems. Therefore, the space propulsion system selection from fuel and oxidizer type aspect and achieving the maximum performance and minimum cost has been performed. Then, the fuel and oxidizer pair samples based on the mass – energy specifications (engine weight- specific impulse) and engine operation cycle type with respect to the mission possibility has been determined. To this end, the algorithm for implementing and using the proposed cost estimation model has been designed. In this algorithm, the proposed cost estimation model is developed based on the existing cost estimation relationship and verified by comparing the existing models. Finally, the outputs in the algorithm are cost-performance (specific impulse) graph for the seven fuels and oxidizer pairwise, engine selection based on achieving maximum specific impulse and providing the design space searches for the cost and time optimization in the space projects.

The main purpose of this paper is to design a navigation CubeSat constellation for local positioning system (LPS) that provides the postiong of Iran, based on COTS approach. To this end, the initial processing of subsystems' information was carried out using statistical processing of CubeSats with the same mission. According to the resultant data, common specifications of different subsystems (structure, electric power, attitude determination and control, thermal control, and communication), including required mass, power, volume and price were determined. Furthemore, the payload design was performed based on the specific mission, design matrix requirements, etc. by utilizing similar subsystem class catalogues (called COTS). Also, local positioning and coverage of the satellite constellation for the Iran was analyzed and simulated from the availabile datd in the Satellite Tool Kit (STK). Finally, the CubeSats constellation with specified subsystem (based on COTS approach) has been suggested for the acceptable coverage percentage to determine the local positioning in Iran. ‌‌

Now, the required samples to achieve the specific precision of sensitivity analysis in design are performed based on trial and error methods. The purpose of this paper is to develop for determining the number of the required sample to achieve the specific precision of sensitivity analysis. Thus, in this paper, a new sensitivity analysis method is proposed based on the Progressive Latin hypercube Sampling (PLHS) and the convergence of the analysis results. For this purpose, a PLHS method has been developed. This cystic approach has led to a sensitivity analysis of accuracy, efficiency and speed in a variety of models with a large number of large parameters and large changes. Sensitivity analysis has been performed on the design of a hydrazine monopropellant thruster catalyst bed model as a case study. The results of this study indicate that in the sensitivity analysis based on the PLHS, the minimum population required for sensitivity analysis with specified accuracy can be determined. This leads to lower processing costs in the sensitivity analysis process, especially in complex models.

The aim of this article is to present the methodology for modeling and simulation of start effects in spacecraft or satellite's propulsion system on the fuel sloshing in the tank by pendulum model in microgravity conditions. In other words, the main aim of this paper is pure sloshing study of fuel and ullage gas relative movement, neglecting the role of Propellant Management Device (PMD). To this end, fuel sloshing in tank is performed by utilizing the Fluent software based on pendulum model. Firstly, algorithm inputs are determined (exiting input, fuel and ullage gas volume, loading, dimensional specifications, etc.); then, tank is modeled and designed and, finally, fuel sloshing simulation in micro-gravity conditions is developed. Fuel sloshing modeling and simulation outputs include determining the sloshing damping rate and its value in the simulation at 20 sec, velocity variation contour, velocity direction contour in the tank, and also ullage gas and fuel relative location in 0.2, 0.4, and 1 sec. The accuracy of the obtained results has been evaluated with the similar experimental results.

The purpose of this article is to optimal manned space launch system conceptual design methodology with combination of modular design (by using clustering the existing motors to provide the thrust force) and sensitivity analysis (by varying the affected parameters to achieve the capability) approach. This methodology is implemented according to the human departure to space program and higher strategy document (country’s aerospace development comprehensive document). To this end, in the methodology is utilized both of the statistical and parametric (the space launch system optimized main parameters) methodologies, is determined the optimum thrust level based on the three fundamental requirements (number of stages, number of engines in clustering and maximum axial acceleration). These fundamental requirements are affected on risk and manned space launch system axial acceleration. In the paper, the purpose of sensitivity analysis is to determine of value of effective of main design parameters on space launch system capabilities. The method for optimizing and design space searching is utilized from Genetic Algorithm (GA). Finally, the suggested methodology and mass – energy capabilities will be verified by comparing the results of two methodology (statistical and optimal) to achieve the specific mission.

The main goal of this paper is to introduce the Moon exploration mission design based on existing technology.The Moon exploration mission design entailsoptimal maneuvering orbit, payload and launch vehicle design. Optimal maneuvering orbit is designed with respect to Circular Restricted Three Body Problem (CRTBP) to model the motion of a spacecraft in the Earth/Moon system. To this end, optimal maneuvering orbitadopted CRTBP as dynamical model and obtained three-dimensional Earth to Moon transfers with low cost. This method is more preferable and flexible than Hohmann transfer because of its lower cost and its access to various inclinations in departure and arrival.The optimal Launch Vehicle Conceptual Design (LVCD) algorithm is based on optimization of major design parameters. LVCD algorithm is coded in a software to let the design engineer explore the design space and to reduce the cost and time of the conceptual design phase that is developed by the authors.The optimization process is performed subject to the restrictions and the performance index is optimized in a mutual iteration mechanism. Consequently, the designed launch vehicle ability to satisfy the mission objectives and its requirements is evaluated.

The major purpose of this paper is to consider the strategy and long term approaches in th American U.S. Space Launch Systems (SLSs)’s technology development. Recently, new approaches has been interested by U.S.. The major roots of these approaches can be searched in the Small Unit Space Transportation and Insertion (SUSTAIN) program that is suggested by U.S. marine corps. The main objective of SUSTAIN is to convey facilities and equipmqnt to the any point of the earth in the minimum time and cost. Hence, the importance of this program leaded to develop the future SLSs’s technology in SUSTAIN orientation. Then, in this paper, the SUSTAIN’s needs is considered and also presented the new approaches in technology development of SLSs to meet the programm’s requirements.

The main aim of this paper is to analysis of chemical performance of hydrogen peroxide based on numerical and parametric methods. The proper chemical function of the catalytic bed, as one of the components of monopropellant thruster, plays a significant role in achieving the two design main goals in (minimizing mass and maximizing the specific impulse). To this end, the effect of catalyst diameter (granules) on the bed chemical performance, optimal length and pressure drop, simulations for beds with different catalytic pellet diameters have been made to 0.4-0.9 cm diameters. Hydrogen peroxide with a concentration of 90% is defined as an inlet fluid at 0.014 m/s in simulations. The calculation of flow pressure drop across the catalyst bed is one of the activities undertaken in this study. The results of this study indicate that with increasing the pellet diameter, the reaction effective surface is reduced and the catalyst bed length is increased for complete decomposition of the propellant. In addition to the required length for complete decomposition of hydrogen peroxide, the pressure drop in various catalyst beds have also been calculated and evaluated. The results of the catalytic bed drop evaluation indicate that at a specific flow rate, a minimum pressure drop will be made in a specific diameter. The reason for this is the interaction of reaction surface and catalyst bed lengths on the pressure drop generated during the propellant decomposition process. Verification and validation of achieved results was conducted by comparing with experimental results.

The purpose of this article is to system design methodology of Propellant Management Device (PMD) for hydrazine fuel tank which used in low (zero) gravity conditions. To this end, the suggestion system design flowchart has three main steps that concluded: step one, Tank design and modeling; step two, PMD design and modeling and step three, stored fuel treatment simulation and analysis. In the design flowchart has performed the result of each step based on mission inputs. Therefore, rejected results in each step led to vary the related parameters. Thus Solid Works software is used to primary PMD and tank modeling. Then, numerical simulation is performed to consider PMD's performance and to illustrate the capillary phenomenon for continues fuel transferring in zero-gravity conditions.Also, numerical methods are used to analysis of the tank and the fuel behavior inside the tank with PMD to optimize system design parameters. Hence, Ansys software used to finalize modelling, analysis, meshing and consideration of fuel behavior in PMD by utilizing the Volume Of Fluid (VOF) method. The optimal system parameters related to specifications of PMD with maximum performance of mass and volume flow rates in zero gravity. In conclusion, by comparing the results (PMD performance) with experimental and existing results will be verified.

The major purpose of this paper is presentation of the technology development phase (modernization or technology management) in space systems and also the review of technology development in space launch system of propulsion system. To this end, firstly, the technology areas in both propulsion systems and space launch systems and technology development are introduced. In this study, technology development and achievement trend in all propulsion subsystems are common. Propulsion system used the technology of kerosene and Nitric Acid in the first the step. Then the Hydrazine fuel, N2O4, L(O2), Kerosene and finally L(O2) and (H2) are achieved. As a conclusion, a new SLS in space transportation is considered and the reason for technology development is clarified so as to find a reduced cost (in both space launch system technology development approaches) and enhance the performance (thrust chamber low cost concepts and reduction of engine components in Expandable Launch Vehicle (ELV) and thrust chamber enhanced life concepts and reliability improvemments in Reusable Launch Vehicle (RLV)).

The purpose of this article is to design the optimum method of Propellant Management Device (PMD) of hydrazine fuel tank which used in zero-gravity conditions. To this end, numerical methods are used to analysis of the tank and the fuel behavior inside the tank with PMD to optimize system design parameters. Hence, Ansys version 17 software used to finalize modelling, analysis, meshing and consideration of fuel behavior in PMD by utilizing the Volume Of Fluid (VOF) method. Also Solid Works software version 2016 is used to primary PMD and tank modeling. Then, numerical simulation is performed to consider PMD's performance and to illustrate the capillary phenomenon for continues fuel transferring in zero-gravity conditions. The design variables in tank and PMD optimization respectively are: minimizing the tank weight to safety factor ratio; dimensional specifications of tank and PMD (height, diameter, length and width dimensions). The objectives of PMD optimization are to achieve maximum volumetric and mass flow rate values. On the other hands, to achieve the most desirable amount of fuel to PMD, that at the end of the time of simulation used by flow rates curves. Numerical analysis results that are obtained include: optimal system parameters related to the specifications of the tank with minimum weight and maximum safety factor and also optimal system parameters related to specifications of PMD with maximum performance of mass and volume flow rates in zero gravity. In conclusion, by comparing the existing systems with the optimal system parameters results will be verified.

The major purpose of this paper is to present Space Launch System (SLS) family technology development from propulsion system aspect. Thus, the models of cost estimation for two types of propulsion systems (cryogenic and semi-cryogenic) are derived based on the statistical method and are then compared with each other. The SLS family modernization model includes five main steps:(1) SLS family propulsion system mass and energetic calculations; (2) Cost estimation and analysis; (3) Sensitivity analysis of propellant volume tanks; (4) Sensitivity analysis of propulsion system performance based on cost; (5) mass, energetic and cost calculations of cryogenic and semi-cryogenic propulsion systems. Finally, the results of the modernization methodology execution are verified by an existing propulsion system.