Propulsion
Mohammad Hossein mansouri Moghari; Hassan Naseh; Sahar Noori
Abstract
Accurate solving of complex systems such as spacecraft is very costly and time consuming. By 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 ...
Read More
Accurate solving of complex systems such as spacecraft is very costly and time consuming. By 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 highly valuable in solving complex systems. The purpose of this article 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
Propulsion
masumeh kiantaj; morteza farhid; mohammadmehdi shafie; Mohammad reza morad
Abstract
The this article the characteristics of the hollow cathode plasma particles in the spt-100 hall effect thruster have been investigated by two-dimensional particle-in-cell simulation. One of the main and important components of the hall thruster is hollow cathode which plays two important tasks: one part ...
Read More
The this article the characteristics of the hollow cathode plasma particles in the spt-100 hall effect thruster have been investigated by two-dimensional particle-in-cell simulation. One of the main and important components of the hall thruster is hollow cathode which plays two important tasks: one part of the electrons that come from the cathode used for anode propellant ionization, and the other part plays an important role of neutralizing the ion beam coming out of the thruster. Therefore, the study of the hollow cathode characteristic is importance. Krypton is used as fuel in this system. Potential changes, density of electrons, ions and temperature of particles have been studied throughout the simulation area. The results show that corresponding to the electrons, the ion density also decreases from the maximum value in the cathode ionization region exponentially through outer chamber. Also, analyzing normalized radius regard to electron density shows that the cathode effective area in which the radius electron temperature reaches maximum value is located about 1.5mm from the center line of the hallow cathode
Propulsion
Nematollah Fouladi; Alireza Mohammadi
Abstract
The purpose of this research is to evaluate a ground test bed of an orbital transmission engine with pre-evacuation of the engine's internal space. In the usual tests on the ground, the initial pressure of the engine is atmospheric pressure. While during the orbital mission, the internal space of the ...
Read More
The purpose of this research is to evaluate a ground test bed of an orbital transmission engine with pre-evacuation of the engine's internal space. In the usual tests on the ground, the initial pressure of the engine is atmospheric pressure. While during the orbital mission, the internal space of the engine may be in the vacuum pressure. Therefore, to ensure the proper performance of internal ballistics, it is necessary to test the performance by pre-evacuating its internal space. In this research, first, the suitability of an exhaust diffuser for this type of test is investigated numerically. Then, the unsteady numerical simulations have been done by applying the pressure-time profiles of the engine as the boundary condition of the inlet pressure. Investigations show that the two phenomena of flow being supersonic in the diffuser at very low engine pressures and the discharge of the return flow to the vacuum chamber prevent the significant influence of environmental conditions on the flow inside the nozzle. So, from the initial moment to the stable working of the diffuser, the flow in the first half of the nozzle is in the supersonic state. Therefore, the internal ballistics of the engine is evaluated independently of the conditions of the outside environment
Propulsion
Nooredin Ghadiri Massoom; alireza rajabi; mohamad ali amirifar; Zahra Amirsardari; Akram Dourani; majid kamranifar
Abstract
In this paper, the effects of different weight percentage of iridium (Ir) nanoparticles loadings on performance parameters of hydrazine catalyst and monopropellant thruster have been studied. Nanoparticles of iridium with different contents of 10 wt%, 20 wt%, and 30 wt% has been coated on gamma-alumina ...
Read More
In this paper, the effects of different weight percentage of iridium (Ir) nanoparticles loadings on performance parameters of hydrazine catalyst and monopropellant thruster have been studied. Nanoparticles of iridium with different contents of 10 wt%, 20 wt%, and 30 wt% has been coated on gamma-alumina of 1 to 2 mm size for decomposition of hydrazine during some various steps of calcination. These catalysts then have been tested in a 1 N thruster. The tests were conducted using a scenario of different stages of steady and pulsating fires of different times and duty cycles. The test results showed that catalyst loss was minimum with 30 wt% of iridium nanoparticles loading. Despite of this, there were no meaningful difference between other parameters such as pressure roughness, thrust, specific impulse, and catalyst crushing. The results showed a good value of characteristic velocity. All parameter values of three type of catalysts were in the expected and desired range
Propulsion
ALI RASTGAR; Hojat Ghasemi
Abstract
The purpose of this paper is to present experimental data for merging two outlet fluid skirts in a dual pressure- swirl coaxial injector. In this study, a dual pressure- swirl external mixing injector was designed and fabricated. Operational characteristics including discharge coefficient and breakup ...
Read More
The purpose of this paper is to present experimental data for merging two outlet fluid skirts in a dual pressure- swirl coaxial injector. In this study, a dual pressure- swirl external mixing injector was designed and fabricated. Operational characteristics including discharge coefficient and breakup length were expressed in terms of different injection pressures for internal and external injectors. Utilizing fast shooting, based on the backlighting method, the interaction between of two outlet skirts were investigated and the merge performance map was extracted. The merge performance map results indicated, when the pressure difference of the external injector increases from 0.3 bar to 0.95 bar, the pressure difference of the internal injector for the merge to occur increases. The reason for this increase in this range of external injector injection pressure differences is that, the effect of the internal injector injection pressure for merge in this area is greater than the external injector injection pressure, the external injector skirt is pulled toward the internal injector skirt. For injection pressures difference of more than 0.95 bar in the external injector, because the effect of the external injector injection pressure for merge is greater than the internal injector injection pressure, the internal injector injection pressure difference is reduced for the merge to occur and the internal injector skirt is pulled toward the external injector skirt.
Propulsion
Hadiseh Karimaei
Abstract
The monopropellant thrusters of the situation control system are a requirement for the development and application of satellites and space capsules in space, which are high-tech and expensive. In this paper, the design and simulation of a pressure- swirl injector with full-cone spray as a fuel injector ...
Read More
The monopropellant thrusters of the situation control system are a requirement for the development and application of satellites and space capsules in space, which are high-tech and expensive. In this paper, the design and simulation of a pressure- swirl injector with full-cone spray as a fuel injector of a monopropellant thruster are presented. For this injector, internal flow simulation was performed in order to predict its output flow characteristics including spray cone angle, output velocity distribution, mass flow rate, spray pattern, etc. For this purpose, VOF fluid volume method is used and the flow turbulence is simulated using the k-eps model. This type of injector is actually a combination of straight flow injector and swirl flow injector. Jet straight flow in the center of the injector and swirl flow along the injector wall are flowed. Both flow regimes are combined in the swirl chamber and the spray is formed as a full-cone. If the ratio of the outlets is selected correctly, the radial and environmental distribution of the liquid jet will be uniform. This injector is preferred to the capillary type (straight flow) and the swirl type. The pressure-swirl injector spray angle is larger than the capillary type, which improves the coverage of the catalyst bed, at the same time, spray angle is not as large as the swirl injector, which enlarges the radial dimensions of the decomposition chamber. Based on the results, it was ensured that the injector provides the desired mass flow rate (about 5.8 gr/s) at a certain design pressure difference (3 bar) and determines a suitable spray pattern. It also provides the desired spray angle (about 35).
Propulsion
mohamad ali amirifar; alireza rajabi; nooredin ghadiri masoom; zahra amirsardari; majid kamranifar
Abstract
In this research, the performance of a monopropellant hydrazine thruster in atmospheric conditions has been investigated experimentally. For this purpose, after designing and constructing the thruster according to the functional requirements of the thruster, a test was designed and after that, the desired ...
Read More
In this research, the performance of a monopropellant hydrazine thruster in atmospheric conditions has been investigated experimentally. For this purpose, after designing and constructing the thruster according to the functional requirements of the thruster, a test was designed and after that, the desired thruster was tested in atmospheric conditions. The test results show that the tested thruster can generate 2000 pulses with a width of 0.5 seconds and a periodicity of one second with reproducibility. It was shown that the life of this thruster is more than 2000 pulses and the thruster was able to produce very small beats of 3 mNS in reproducibility. Also, comparing the results of the current thruster sample with the experimental results of other thrusters showed how by selecting the appropriate dimensions for the injector, catalyst chamber and nozzle, the characteristics of pressure rise time, minimum impulse, pulse centroid and pressure drop time in the Thruster can be well controlled. Reducing the injector diameter (by keeping the flow rate constant by increasing the injection pressure) reduces the impulse (within a constant pulse width) and increases the pressure rise time. Reducing the dimensions of the catalyst chamber also reduces the increase and decrease time of the pressure, resulting in a smaller pulse centroid
Propulsion
Nematollah Fouladi; sina afkhami; Mahmood PasandidehFard
Abstract
In the present study, the effect of pre- evacuation on starting process of a second throat exhaust diffuser has been investigated experimentally by examining a thrust optimized parabolic nozzle. An experimental setup called high- altitude test facility has been used with compressed air as operating fluid. ...
Read More
In the present study, the effect of pre- evacuation on starting process of a second throat exhaust diffuser has been investigated experimentally by examining a thrust optimized parabolic nozzle. An experimental setup called high- altitude test facility has been used with compressed air as operating fluid. According to the importance of area ratio parameter (Ad/Ast) of a second throat diffuser, the effect of this parameter variation has been examined on the start- up performance of the nozzle and diffuser. In each of the diffuser geometries, in order to evaluate the instantaneous performances, the pressure in the nozzle chamber has charged instantly in two modes; with and without pre- evacuation. Then, the vacuum chamber pressure and static pressure distribution along the diffuser were measured by a data acquisition system. The results show that pre- evacuation in the test chamber reduces the start- up time of the diffuser by 50 to 60%. In addition, pre- evacuating the test chamber eliminates the destructive transition phenomenon from the flow separation pattern during start- up of the nozzle and diffuser. Also, It has been observed that with the narrowing of the diffuser’s second throat duct, the minimum starting pressure of the diffuser increases and eventually flow chocks at the second throat in a certain area ratio.
Propulsion
seyed alireza Jalali chimeh; Ali Madadi; seyed mostafa safavi homami; javad Emami
Abstract
In complicated systems such as satellites, each subsystem's design can affect the design of the overall system. In the design procedure, the effect of each technology on the other components should be determined. Because of increasing in space trash, the propulsion subsystems are introduced to avoid ...
Read More
In complicated systems such as satellites, each subsystem's design can affect the design of the overall system. In the design procedure, the effect of each technology on the other components should be determined. Because of increasing in space trash, the propulsion subsystems are introduced to avoid collision in space. One of the methods to attain high altitude orbits is flight maneuvers using propulsion systems. Several types of propulsion systems are utilized in satellites. Resistojets can be employed as a low-cost propulsion systems for satellites because they do not use complicated technologies. In the present research, a resistojet propulsion system is designed for a CubeSat for the mission of orbital altitude reduction. The propellant is selected according to all properties. The design of the nozzle and the heater is also introduced, The overall layout of the system is presented and finally, an algorithm of electrical propulsion systems for a specified mission is proposed.
Propulsion
Sahar noori; Rojin Shokri Khanghah; mohammad nadafipour meibody
Abstract
Microwave electrothermal thruster is the propellant that converts microwave energy into heat energy. nowadays, with the technology development of electric thrusters are very important in terms of producing high specific impact and low fuel consumption. These thrusters can produce acceptable thrust over ...
Read More
Microwave electrothermal thruster is the propellant that converts microwave energy into heat energy. nowadays, with the technology development of electric thrusters are very important in terms of producing high specific impact and low fuel consumption. These thrusters can produce acceptable thrust over a long period of time, which are suitable for spiral orbital transfer missions. In this type of propulsion, the propellant gas is heated and expanded, which produces the propulsion force. This paper presents the conceptual design process of a microwave thermal thruster. The propulsion system includes propellant, propulsion storage tank, amplification chamber, and power plant, which includes batteries and solar arrays. In this paper, the method of calculating the mass and the characteristics of each are presented in detail. Finally, in order to validate the conceptual design process presented in this study, the necessary studies have been discussed. Conceptual design has been done for a 100 kg satellite, which is desirable to travel in a week from an orbital height of 300 to 800 km during a spiral treansfer. The propulsion system and mass of each subsystem are obtained.
