Computational Study on a Single Flow Element in a Nuclear Thermal Rocket

2013 ◽  
Author(s):  
Xiang Zhao ◽  
Trent Montgomery ◽  
Sijun Zhang
Keyword(s):  
Computational Study ◽  
Specific Impulse ◽  
Propulsion Systems ◽  
Future Space ◽  
Nuclear Propulsion ◽  
Stress Estimation ◽  
Flow Through ◽  
Single Flow ◽  
Chemical Propulsion ◽  
Flow Element

The nuclear thermal rocket is one of the candidate propulsion systems for future space exploration including traveling to Mars and other planets of the solar system. Nuclear thermal propulsion can provide a much higher specific impulse than the best chemical propulsion available today. A basic nuclear propulsion system consists of one or several nuclear reactors that heat hydrogen propellant to high temperatures and then allow the heated hydrogen and its reacting product to flow through a nozzle to produce thrust. This paper presents computational study on a single flow element in a nuclear thermal rocket. The computational results provide both detailed and global thermo-fluid environments of a single flow element for thermal stress estimation and insight for possible occurrence of mid-section corrosion.

Micropropulsion Development at the ARC Seibersdorf Research

2006 ◽  
Author(s):  
Carsten Arthur Scharlemann ◽  
Martin Tajmar
Keyword(s):  
Hydrogen Peroxide ◽  
Electric Propulsion ◽  
Specific Impulse ◽  
Electrical Power ◽  
Preliminary Test ◽  
Small Mass ◽  
Space Missions ◽  
Propulsion Systems ◽  
Future Space ◽  
Chemical Propulsion

The increasing application of micro-satellites (from 10 kg up to 100 kg) for a rising number of various missions demands the development of new miniaturized propulsion systems. Micro-satellites have special requirements for the propulsion system such as small mass, reduced volume, and very stringent electrical power constraints. Existing propulsion systems often can not satisfy these requirements. The Space Propulsion Department of the ARC Seibersdorf research dedicated itself to the development and test of various micropropulsion systems for present and future space missions. The portfolio of the systems under development includes electrical and chemical propulsion systems. The covered thrust and specific impulse of the developed propulsion systems ranges from 1μN to 1N and 500 s to 8000 s respectively. Based on the large experience obtained over several decades in the development of Field Emission Electric Propulsion systems (FEEP), several microstructured FEEPs have been developed. The design of these systems is presented as well as preliminary test results and a summarization of the experience obtained during the process of miniaturizing such systems. The development of miniaturized chemical propulsion systems includes a bipropellant and a monopropellant thruster. The bipropellant thruster constitutes the smallest existing 1N thruster utilizing hydrogen peroxide. The thruster system consists of two micopumps for the propellant feed and a microturbine to generate the power for operating the pumps. The monopropellant thruster is a derivative of the bipropellant thruster. It offers a lower specific impulse than the bipropellant system but due to its reduced system complexity it represents also a promising candidate for several future space missions. Both systems utilize rocket grade hydrogen peroxide (green propellant), which is decomposed with the help of an advanced monolithic catalyst. The present paper discusses the design methods and the physical limitations of such chemical propulsion systems with regard to their miniaturization and summarizes their performance evaluation.

scholarly journals MPD Thruster Technology and Its Limitations

2021 ◽  
Vol 2 (4) ◽  
Author(s):  
Samarth Patel ◽  
M.S.R. Bondugula ◽  
Srilochan Gorakula
Keyword(s):  
Electric Propulsion ◽  
Specific Impulse ◽  
Empirical Models ◽  
Scaling Factor ◽  
Electrode Erosion ◽  
Propulsion Systems ◽  
Chemical Propulsion ◽  
Mpd Thrusters ◽  
Limiting Value ◽  
Electromagnetic Propulsion

It was realized earlier that chemical propulsion systems utilize fuel very inefficiently, which greatly limits their lifespan. Electric propulsion is into existence to overcome this limitation of chemical propulsion. The magnetoplasmadynamic (MPD) thruster is presently the most powerful form of electromagnetic propulsion. It is the thruster’s ability to efficiently convert MW of electric power into thrust which gives this technology a potential to perform several orbital as well as deep space missions. MPD thruster offers distinct advantages over conventional types of propulsion for several mission applications with its high specific impulse and exhaust velocities. However, MPD thruster has limitations which limits its operational efficiency and lifetime. In this paper, the thruster limitations are reviewed with respect to three operational limits i.e., the onset phenomenon, cathode lifetime, and thruster overfed limits. The dependence and effects of the operational limits on each other is compared using different empirical models to derive a scaling factor that has been found for each geometrical arrangement; a limiting value exists beyond which the operation becomes highly unsteady and electrode erosion occurs. Along with reviewing and proposing methods to overcome power limitations for MPD thrusters, the relation between exit velocity and ratio of electrode’s radius is also verified using Maecker’s formula.

Investigation of Variation in the Performance of an Electro Thermal Thruster with Aerospike Nozzle

2016 ◽  
Vol 16 ◽  
pp. 91-103
Author(s):  
Pranav Menon
Keyword(s):  
Electric Propulsion ◽  
Specific Impulse ◽  
Velocity Variation ◽  
Propulsion Systems ◽  
Exit Velocity ◽  
Hot Plasma ◽  
Exit Nozzle ◽  
Ion Propulsion ◽  
Aerospike Nozzle ◽  
Chemical Propulsion

One of the most recently developed modes of propulsion is electric propulsion. The commonly used chemical propulsion systems have the advantage of a high Specific Impulse as compared to that of ion propulsion systems. However, owing to the efficacy of ion propulsion systems, it is considered the future of space exploration.Electro thermal thrusters produce thrust by using electrical fields to force hot plasma out of the nozzle with certain exit velocity. The plasma’s exit velocity and the system’s thrust capacity, as of now, are insufficient for space travel to be conducted within a reasonable time. I intend to study the possibility of improving the thruster’s performance by using an aerospike nozzle as an exit nozzle which meets the conditions required for the thruster to function appropriately. I shall be studying the plasma plume exit velocity variation with respect to the nozzles used. Also, a thermal analysis will be conducted in order to find the correct material for the nozzle.

Simulation of Cooling Systems in Gas Turbines

1996 ◽  
Vol 118 (2) ◽  
pp. 301-306 ◽  
Author(s):  
G. Ebenhoch ◽  
T. M. Speer
Keyword(s):  
Gas Turbines ◽  
Path Following ◽  
Cooling Systems ◽  
Turbine Engine ◽  
Rotating Systems ◽  
Entire Flow ◽  
Flow Pressure ◽  
The One ◽  
Single Flow ◽  
Flow Element

The design of cooling systems for gas turbine engine blades and vanes calls for efficient simulation programs. The main purpose of the described program is to determine the complete boundary condition at the coolant side to support a temperature calculation for the solid. For the simulation of convection and heat pick up of the coolant flow, pressure loss, and further effects to be found in a rotating frame, the cooling systems are represented by networks of nodes and flow elements. Within each flow element the fluid flow is modeled by a system of ordinary differential equations based on the one-dimensional conservation of mass, momentum, and energy. In this respect, the computer program differs from many other network computation programs. Concerning cooling configurations in rotating systems, the solution for a single flow element or the entire flow system is not guaranteed to be unique. This is due to rotational forces in combination with heat transfer and causes considerable computational difficulties, which can be overcome by a special path following method in which the angular velocity is selected as the parameter of homotopy. Results of the program are compared with measurements for three applications.

Simultaneous determination of thiamine and pyridoxine in pharmaceuticals by using a single flow-through biparameter sensor

2001 ◽  
Vol 25 (3-4) ◽  
pp. 619-630 ◽  
Author(s):  
P Ortega Barrales ◽  
A Domı́nguez Vidal ◽  
M.L Fernández de Córdova ◽  
A Molina Dı́az
Keyword(s):  
Simultaneous Determination ◽  
Flow Through ◽  
Single Flow

An analytical and experimental study of a hybrid rocket motor

2014 ◽  
Vol 228 (13) ◽  
pp. 2475-2486 ◽  
Author(s):  
Hadi Rezaei ◽  
Mohammad Reza Soltani
Keyword(s):  
Flow Rate ◽  
Low Cost ◽  
Specific Impulse ◽  
Combustion Products ◽  
Rocket Motor ◽  
Chamber Pressure ◽  
Hybrid Rocket ◽  
Hybrid Rocket Motor ◽  
Chemical Propulsion ◽  
Rate Case

The hybrid rocket motor is a kind of chemical propulsion system that has been recently given serious attention by various industries and research centers. The relative simplicity, safety and low cost of this motor, in comparison with other chemical propulsion motors, are the most important reasons for such interest. Moreover, throttle-ability and thrust variability on demand are additional advantages of this type of motor. In this paper, the result of an internal ballistic simulation of hybrid rocket motor in a zero-dimensional form is presented. Further to validate the code, an experimental setup was designed and manufactured. The simulation results are compared with the experimental data and good agreement is achieved. The effect of various parameters on the motor performance and on the combustion products is also investigated. It is found that increasing the oxidizer flow rate, increases the pressure and specific impulse of the motor; however, the slope of the specific impulse for the high flow rate case reduces. In addition, by increasing the combustion chamber pressure, the specific impulse is increased considerably. The initial diameter of the fuel port does not have significant effect on the pressure chamber and on the specific impulse. Addition of a percentage of an oxidizer like ammonium perchlorate to the fuel increases the specific impulse linearly.

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