Space systems design (spacecraft, satellites, space stations and their equipment)
Sajjad Gharezi; Mohammad Mehdi Doustdar
Articles in Press, Accepted Manuscript, Available Online from 17 January 2024
Abstract
In this research, the design of a can combustion chamber for a ramjet engine, the performance of this chamber and the role of flameholder have been studied. For this purpose, after talking about ramjet engine types, some general information about the combustion chamber has been explained. Then the process ...
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In this research, the design of a can combustion chamber for a ramjet engine, the performance of this chamber and the role of flameholder have been studied. For this purpose, after talking about ramjet engine types, some general information about the combustion chamber has been explained. Then the process of chamber designing and determining its geometry has been discussed. The dimensions of the chamber were determined by using input conditions extracted from GasTurb software as well as applying a calculation code, and the geometry of the chamber have been determined by applying calculation codes. By evaluating the obtained geometry and ensuring the accuracy of the design, simulation of combustion with non-premixed liquid phase was carried out using Fluent software. While presenting the results, the effects of size, distance and number of flameholder have been investigated. It is shown that the use of flameholder in ramjet engines is essential but the use of large flameholder is not recommended.
Space systems design (spacecraft, satellites, space stations and their equipment)
Vahid Rahimi Ghoradel; Hossein Mahdavy-Moghaddam
Volume 16, Issue 2 , June 2023, , Pages 1-17
Abstract
Each missile has a payload section and an engine section. In the path of the missile, there is a time when the mission of the engine section is over and after that the engine will not play an effective role and will be as extra weight and consequently reduced range or factor for easy detection of the ...
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Each missile has a payload section and an engine section. In the path of the missile, there is a time when the mission of the engine section is over and after that the engine will not play an effective role and will be as extra weight and consequently reduced range or factor for easy detection of the warhead by enemy agents. In such a situation, after completing the engine mission, the mechanism of separating the steps and separating the head from the body is used. One of the separation methods is to use the thrust termination system method. In this paper, with the studies performed on the thrust termination system and the presentation of mathematical relations, the pressure drop and inverse thrust created in the chamber after opening the reverse thrust valves are predicted. Also, cold type separation and thrust termination system were used and the combustion chamber pressure drop is simulated. Then, the effect of important and influential factors on the thrust termination system has been investigated.
Space systems design (spacecraft, satellites, space stations and their equipment)
Sajjad Davari; Hadiseh Karimaei; Mohammad Reza Salimi; Hassan Naseh
Volume 16, Issue 2 , June 2023, , Pages 55-61
Abstract
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 ...
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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.
Space systems design (spacecraft, satellites, space stations and their equipment)
Sajjad Davari; Hadiseh Karimaei; Mohammad Reza Salimi; Hassan Naseh
Volume 16, Issue 1 , March 2023, , Pages 35-46
Abstract
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 ...
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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.
Space systems design (spacecraft, satellites, space stations and their equipment)
Hojat Ghasemi; Seyed Mohammadreza Mahmoudian; Noordin Qadiri Massoom; S. Rashad Rouholamini; Pouria Mikaniki; Asghar Azimi
Volume 16, Issue 1 , March 2023, , Pages 47-58
Abstract
The aim of the present research is to obtain the ability to use the cryogenic propellant engines on a laboratory scale. In this regard, it is necessary to build some experimental motors and investigate the their performance parameters. The liquid oxygen as a common oxidizer and ethanol as a green fuel ...
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The aim of the present research is to obtain the ability to use the cryogenic propellant engines on a laboratory scale. In this regard, it is necessary to build some experimental motors and investigate the their performance parameters. The liquid oxygen as a common oxidizer and ethanol as a green fuel have been selected as propellant components. The engine is designed to produce 400 kgf force at the nominal condition. The pintle type injector has been chosen in which liquid oxygen and fuel are flowed in the axial and radial directions, respectively. The combustion chamber has been protected against overheating by applying the regenerative cooling. However, the laboratory feature of the engine design has provided the using of water instead the cooling propellant. All main components of the engine such as injector, igniter, and flow controllers, are examined by the cold tests. A comprehensive test facility is designed and set up for hot fire tests in which the performance of almost all parameters can be evaluated. Fifteen fire tests have been performed. Maximum obtained pressure and evaluated combustion efficiency were about 75% of design values.
Space systems design (spacecraft, satellites, space stations and their equipment)
Amirhamzeh Farajollahi; Reza Firuzi; Mohammad Reza Salimi; Mohsen Rostami
Volume 15, Issue 4 , December 2022, , Pages 107-121
Abstract
In this study, the effects of geometry and spiral rifling like guides inside the injection nozzle on the performance of an engine are investigated, using AVL Fire software. To do so, firstly injectors with different nozzle geometries and their resultant spray patterns were simulated. Numerical results ...
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In this study, the effects of geometry and spiral rifling like guides inside the injection nozzle on the performance of an engine are investigated, using AVL Fire software. To do so, firstly injectors with different nozzle geometries and their resultant spray patterns were simulated. Numerical results of this step show that creation of spiral rifling like guides inside the nozzle increases the spray cone angle and improves fuel atomization quality. In the next step, effects of using forgoing nozzle geometries on sample engine characteristics were studied and the related results compared to those of common cylindrical injectors. Numerical results of this step clearly show the superior performance of nozzles with spiral rifling like guides. In this case, SFC reduces up to 32 percent while the engine power and it's torque rises more than 63 percent. Also the amount of pollutants like NOx reduces 12 percent with respect to common cylindrical nozzles.
Space systems design (spacecraft, satellites, space stations and their equipment)
Sajad Davari; Hadiseh Karimaei
Volume 15, Issue 3 , September 2022, , Pages 109-118
Abstract
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 ...
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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.
Space systems design (spacecraft, satellites, space stations and their equipment)
Hamed Alisadeghi; Azade Khadivi; Ehsan Zabihian
Volume 15, Issue 2 , June 2022, , Pages 43-58
Abstract
The interaction of thruster plumes with satellite components can have undesirable effects, such as disturbance force/torque, thermal loading, and species deposition in the surfaces. The purpose of this paper is to use the Direct Simulation Monte Carlo (DSMC) method to analyze the 3D plume impingement ...
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The interaction of thruster plumes with satellite components can have undesirable effects, such as disturbance force/torque, thermal loading, and species deposition in the surfaces. The purpose of this paper is to use the Direct Simulation Monte Carlo (DSMC) method to analyze the 3D plume impingement flows and investigate its effects. Two impingement problems are computed. The impact of a jet of nitrogen on an inclined flat plate is considered. Good agreement is found between surface quantities calculated by DSMC and experimental data. The plume of a hydrazine control thruster firing in a model satellite configuration is simulated. Surface quantities and net impingement effects are calculated. The effects of partial displacement of the thruster locations on the results have also been investigated. The results show that a 20% displacement of the thruster location can change the disturbance force/torque by up to 15% of the initial values.
Space systems design (spacecraft, satellites, space stations and their equipment)
Hamed Moeini; Ebrahim Afshari; Karim Mazaheri
Volume 15, English Special Issue , May 2022, , Pages 1-13
Abstract
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 ...
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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.
Space systems design (spacecraft, satellites, space stations and their equipment)
Ramin Kamali Moghadam; Mohammad Taeibi Rahni; Salar Heyat Davoudian; Reinhard Miller
Volume 15, English Special Issue , May 2022, , Pages 25-33
Abstract
Superhydrophobic coatings can be made by creating a micro-sized structure on a surface providing super-repellent properties which has many applications in aerospace, defense, automotive, biomedical and engineering. Numerical simulation of drop dynamics and motion on a superhydrophobic surface helps us ...
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Superhydrophobic coatings can be made by creating a micro-sized structure on a surface providing super-repellent properties which has many applications in aerospace, defense, automotive, biomedical and engineering. 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. Effect of microstructures on droplet behavior on a superhydrophobic surface is investigated using different microstructures. 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 microstructure surfaces in the same conditions. Comparison of the results shows the effect of microstructure coating on the surface hydrophobicity properties.
Space systems design (spacecraft, satellites, space stations and their equipment)
َAlireza Alikhani; Mohammad Reza Salimi
Volume 15, English Special Issue , May 2022, , Pages 55-64
Abstract
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 ...
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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.
Space systems design (spacecraft, satellites, space stations and their equipment)
Mohammad Reza Salimi
Volume 15, Issue 1 , March 2022, , Pages 89-105
Abstract
In present study, a hydrazine based monopropellant thruster decomposition chamber is simulated numerically. The catalyst bed separated in two sides, the particles size in upstream side is larger than those in downstream side. Effects of upstream side length and its particles diameter on catalyst bed ...
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In present study, a hydrazine based monopropellant thruster decomposition chamber is simulated numerically. The catalyst bed separated in two sides, the particles size in upstream side is larger than those in downstream side. Effects of upstream side length and its particles diameter on catalyst bed characteristics were investigated. To this end, three standard particles sizes of mesh: 16.5, 25 and 30 for the upstream side and two standard particles diameter of 1/8 and 1/16 (in) for downstream side were analyzed. Additionally, three upstream side lengths of 2.5, 5 and 7.5 (mm) were used while the length of bed is 6.5 (cm). Simulations were performed in three bed loading coefficients of 16.5, 25 and 35 (kg/m2s). The related results showed the effectiveness of upstream side on flow and thermal fields are strongly depends on the ration of particles sizes in upstream and downstream sides. Moreover, the upstream side length and bed loading are two important factors affecting the upstream side effectiveness.
Space systems design (spacecraft, satellites, space stations and their equipment)
Hanieh Eshaghnia; Mehran Nosratollahi; Amirhossain Adami; Hadi Dastoury
Volume 15, Issue 1 , March 2022, , Pages 121-137
Abstract
Turbopump propulsion systems have been used in almost all launch vehicles. With the advancement of manufacturing technologies, especially in the use of composite and lightweight structures, the use of non-turbopump propulsion systems has been considered due to the reduction of operating costs. This study ...
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Turbopump propulsion systems have been used in almost all launch vehicles. With the advancement of manufacturing technologies, especially in the use of composite and lightweight structures, the use of non-turbopump propulsion systems has been considered due to the reduction of operating costs. This study has been investigated the multi-disciplinary optimization design of a two-stage launch vehicle using a pressure-fed propulsion system for both stages. Two main propulsion systems including gas-pressure and self-pressure feeding systems, have been evaluated in different configurations on two launcher stages. To extracting the optimum and possible solution, the launcher mission also has been added as a design variable in the optimization algorithm. The launcher has been extracted and introduced for each specific configuration of the launcher to achieve a certain orbital altitude with the maximum carrying payload and minimum gross mass. For this purpose, the AAO multidisciplinary optimization design framework has been used. The system-level and subsystem optimizer of the GA-SQP algorithm have been chosen.
Space systems design (spacecraft, satellites, space stations and their equipment)
Mostafa Jafarpanah; Hassan Naseh
Volume 14, Issue 4 , December 2021, , Pages 25-33
Abstract
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 ...
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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.
Space systems design (spacecraft, satellites, space stations and their equipment)
Hanieh Eshaghnia; Mehran Nosratollahi; Amirhossain Adami
Volume 14, Issue 4 , December 2021, , Pages 35-49
Abstract
A new approach to the design and development of launchers is the use of advanced technologies to reduce design and development costs as much as possible. In this paper, an approach to reduce costs and increase reliability is proposed, which is based on the use of a non-turbo pump propulsion system (pressure-fed ...
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A new approach to the design and development of launchers is the use of advanced technologies to reduce design and development costs as much as possible. In this paper, an approach to reduce costs and increase reliability is proposed, which is based on the use of a non-turbo pump propulsion system (pressure-fed propulsion system) instead of a turbo pump propulsion system. For this purpose, the multidisciplinary conceptual design optimization of a two-stage launch vehicle with a pressure-fed propulsion system with the aim of sending max payload with a least gross mass to the orbit (500 km) in terms of structure, aerodynamics, propulsion, pressure vessels, simulation, and pitch program disciplines. Then, the sensitivity analysis was performed on the optimum launcher to determine the efficiency of the launcher at different orbital heights and the ability to carry a suitable payload.
Space systems design (spacecraft, satellites, space stations and their equipment)
Hojat Taei; Amirhossain Adami; Mansour Hozuri
Volume 14, Issue 4 , December 2021, , Pages 85-98
Abstract
The need to improve the reliability and safety requirements, has led to increasingly utilization of reliability based design approaches. In this study, reliability based multidisciplinary design optimization for a bipropellant propulsion system has been investigated. The objective function is minimizing ...
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The need to improve the reliability and safety requirements, has led to increasingly utilization of reliability based design approaches. In this study, reliability based multidisciplinary design optimization for a bipropellant propulsion system has been investigated. The objective function is minimizing the total system mass and design constraints are the total impulse and the temperature of the wall of the combustion chamber. Monte Carlo simulation methodology is used to apply uncertainties in the problem and to show the reliability of the system under these uncertainties. The mass, functional and geometric results of the bipropellant propulsion system are differentiated for optimal design, reliability based design and optimal reliability based design. Then, considering the results, the concepts and definitions of design methods are compared and discussed and it is shown that the reliability based multidisciplinary optimization while having the desired mass, has high reliability.
Space systems design (spacecraft, satellites, space stations and their equipment)
Hojat Taei; Mahmood Haghighat Esfahani; Sajjad Yadegari Dehkordi
Volume 14, Issue 3 , September 2021, , Pages 39-50
Abstract
In this paper, a novel Comprehensive Preference-based Design (CPD) approach is presented which attempts to achieve subjective attributes that are defined in the concept of maximization of designer/customer's satisfaction in addition to objective goals which are formulated in the form of minimization ...
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In this paper, a novel Comprehensive Preference-based Design (CPD) approach is presented which attempts to achieve subjective attributes that are defined in the concept of maximization of designer/customer's satisfaction in addition to objective goals which are formulated in the form of minimization of a performance criterion in a two-phase structure using two nested optimizers.In the first phase of CPD,using the concept of satisfaction,the subjective preferences of the designer/customer are defined in terms of fuzzy relationships and operators.Whereas the results of this phase are inaccurate,in the second phase,it is attempted to define a performance criterion and in order to achieve an optimal operational plan,attitude parameters and the compromises needed to meet the designer/customer's preferences are implemented.The methodology is utilized to design of a space launch vehicle for delivering 1200 kg payload to a 750 km orbit.Comparison of the results shows that despite the higher mass of launch vehicle designed by CPD,overall design satisfaction is higher and designer/customer's preferences have been satisfied.
Space systems design (spacecraft, satellites, space stations and their equipment)
Mohammad Razmjooei; Mohammad Shahbazi; Fathollah Ommi
Volume 14, Issue 2 , June 2021, , Pages 1-26
Abstract
In this paper, the heat transfer and ablation thermal insulators in solid rocket motor are investigated. Therefore, by collecting and solving the thermal ablation equations, a computer program, using MATLAB software, is developed which can predict the thermal response of insulators in different operating ...
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In this paper, the heat transfer and ablation thermal insulators in solid rocket motor are investigated. Therefore, by collecting and solving the thermal ablation equations, a computer program, using MATLAB software, is developed which can predict the thermal response of insulators in different operating conditions and compare the performance of these insulators. The heat and mass transfer equations are considered in two dimensions in a solid body. We used the equations, finite volume method with implicit formulation for time dependency to solve equations. The reaction equation which written in the form of Arrhenius, is solved using Runge-Kutta method, and the density and the flux of the gas produced at each step are obtained. Also we represent a model for the rate of recession.
Space systems design (spacecraft, satellites, space stations and their equipment)
Masoud EidiAttarZade; Atiyeh SarAbadani; Ghazal Davarnia; Hamed Khosrobeygi; Mohammad Farshchi; Alireza Ramezani
Volume 14, Issue 2 , June 2021, , Pages 47-37
Abstract
Numerical modeling of space engines aside the experimental test is routine. In the design step of such engines, low-cost softwares are vital. In this paper, small-scale space engine thrust chamber analysis code will be developed. In this code, propellant injection and evaporation distribution will be ...
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Numerical modeling of space engines aside the experimental test is routine. In the design step of such engines, low-cost softwares are vital. In this paper, small-scale space engine thrust chamber analysis code will be developed. In this code, propellant injection and evaporation distribution will be modelled. 1D Combustion solver calculates the reactions of distributed fuel and oxidizer through the thrust chamber axis by chemical mechanisms. Then the cooling solver computes the heat transfer from hot gases to the film cooling layer and the outer surroundings. Validation shows acceptable errors in the modelling of processes. By this developed code, the performance of the Astrium bi-propellant thruster with MonoMethylHydrazine and NitrogenTetrOxide and distributed chemical reaction has been investigated. Results show that hot gas temperature inside the combustor is not uniform and has a peak. Furthermore, the evaporation of the propellant droplets is continued to the nozzle throat.
Space systems design (spacecraft, satellites, space stations and their equipment)
Alireza Mohammadi; Fathollah Ommi
Volume 13, Issue 4 , December 2020, , Pages 15-23
Abstract
This paper presents numerical study on spray characteristics and droplet distribution by using Lagrangian method in the discrete phase model of CFD. A two-fluid Eulerian method and Lagrangian approach is selected for modeling two phases turbulence flow in mixing chamber and atomization at outlet of nozzle ...
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This paper presents numerical study on spray characteristics and droplet distribution by using Lagrangian method in the discrete phase model of CFD. A two-fluid Eulerian method and Lagrangian approach is selected for modeling two phases turbulence flow in mixing chamber and atomization at outlet of nozzle while turbulence has been modeled by K-ɛ. In this study, water has been used instead of fuel and Nitrogen instead of atomization gas or oxidizer, while their ratio has been considered 0.32 to provide 26 degrees cone angle and this way, droplet‘s characteristic has been studied and compared with maximum entropy methods. Then droplet‘s diameter has been investigated by changing liquid and gas phase flow rateand based on that, we can optimize atomizer ‘s working condition with maximum efficiency with respect to its cone angle, droplet ‘s diameter and velocity and level of penetration by minimum need of experimental tests.
Space systems design (spacecraft, satellites, space stations and their equipment)
Mehran Nosratollahi; Mohammad Fatehi Fatehi; Amirhossain Adami
Volume 13, Issue 3 , September 2020, , Pages 1-16
Abstract
Orbital transfer blocks has the task of transferring satellites to objective orbits from parking orbit. In this paper, Attention will be given to multidisciplinary optimal design of the propulsion system of two liquid component which is one of the most important subsystems of Orbital transfer blocks. ...
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Orbital transfer blocks has the task of transferring satellites to objective orbits from parking orbit. In this paper, Attention will be given to multidisciplinary optimal design of the propulsion system of two liquid component which is one of the most important subsystems of Orbital transfer blocks. Designing with multi objective bipropellant system, based on minimum total mass and maximum Isp, and at the end mentioned to costs and compared. For combinations of NTO as Oxidizer and fuels which are: UDMH, MMH, Hydrazine and RP-1 then for usual structures that utilized in this systems, design and optimization occurred by multi objective hybrid Particle Swarm Optimization (PSO) algorithms.
Space systems design (spacecraft, satellites, space stations and their equipment)
Ghasem Heydari; Maryam Kiani; S. Hossein Pourtakdost; Mohammad Sayanjali
Volume 13, Issue 3 , September 2020, , Pages 25-38
Abstract
Halo orbits are of importance for observation and study of the space due to their specific characteristics including the orbital position and the periodic motion. In this regards, present paper has focused on optimal trajectory planning to transfer to halo orbits. To this aim, homotopy approach has been ...
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Halo orbits are of importance for observation and study of the space due to their specific characteristics including the orbital position and the periodic motion. In this regards, present paper has focused on optimal trajectory planning to transfer to halo orbits. To this aim, homotopy approach has been adopted for optimal trajectory design. This approach has improved the convergence rate and insensitivity of the problem to initial guess. The designed trajectory transfers a spacecraft orbiting the Earth to a Halo orbit around Lagrangian point L1 of the Earth-moon restricted three-body system. The propulsion system has been assumed to be low thrust with constant specific impulse. Homotopy approach has a broad domain of applicability and methods in which continuation method has been employed here among them. The optimal designed trajectory minimizes the fuel consumption via transforming solution of the minimum energy problem utilizing the homotopy approach. This approach simplifies solution of the complex problem of minimum fuel indeed.
Space systems design (spacecraft, satellites, space stations and their equipment)
mohammad razmjooei; mohammad shahbazi; Fathollah Ommi
Volume 13, Issue 2 , June 2020, , Pages 13-35
Abstract
In this paper, the heat transfer and ablation thermal insulators in solid rocket motor are investigated. Therefore, by collecting and solving the thermal ablation equations, a computer program, using MATLAB software, is developed which can predict the thermal response of insulators in different operating ...
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In this paper, the heat transfer and ablation thermal insulators in solid rocket motor are investigated. Therefore, by collecting and solving the thermal ablation equations, a computer program, using MATLAB software, is developed which can predict the thermal response of insulators in different operating conditions and compare the performance of these insulators. The heat and mass transfer equations are considered in two dimensions in a solid body. We used the equations, finite volume method with implicit formulation for time dependency to solve equations. The reaction equation which written in the form of Arrhenius, is solved using Runge-Kutta method, and the density and the flux of the gas produced at each step are obtained. Also we represent a model for the rate of recession.
Space systems design (spacecraft, satellites, space stations and their equipment)
Hojat Taei; Pourya Shokrolahi
Volume 13, Issue 2 , June 2020, , Pages 87-96
Abstract
The final phase of orbital rendezvous and docking has been studied in this article. The main objective is to control the position of a chaser that can reach the target in the minimum time, or in other words, bypassing the optimal path. Another important objective of this paper is the minimum energy consumption. ...
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The final phase of orbital rendezvous and docking has been studied in this article. The main objective is to control the position of a chaser that can reach the target in the minimum time, or in other words, bypassing the optimal path. Another important objective of this paper is the minimum energy consumption. In the dynamic simulation, the equations of the linear form of Clohessy-Wiltshire (CWH) equations have been utilized. In linear CWH equations, the change in either direction of X or Y will result in the change in another direction and will affect the orbital docking operation. In order to achieve the objectives of this paper, the design variables should be optimized; To optimize the design variables, two methods, i.e. genetic algorithm (GA) and particle swarm optimization (PSO), have been used. Finally, to evaluate the real conditions, the results will be investigated by applying uncertainty in the outputs of thrusters.
