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.

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.

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.

scholarly journals Validation of a Test Platform to Qualify Miniaturized Electric Propulsion Systems

Aerospace ◽  
2019 ◽  
Vol 6 (9) ◽  
pp. 99 ◽  
Author(s):  
Fabrizio Stesina
Keyword(s):  
Electric Propulsion ◽  
Can Bus ◽  
Propulsion System ◽  
Chemical Contamination ◽  
Propulsion Systems ◽  
Electric Propulsion System ◽  
Future Space ◽  
Test Platform ◽  
Electro Magnetic ◽  
Exchange Data

Miniaturized electric propulsion systems are one of the main technologies that could increase interest in CubeSats for future space missions. However, the integration of miniaturized propulsion systems in modern CubeSat platforms presents some issues due to the mutual interactions in terms of power consumption, chemical contamination and generated thermal and electro-magnetic environments. The present paper deals with the validation of a flexible test platform to assess the interaction of propulsion systems with CubeSat-technologies from mechanical, electrical, magnetic, and chemical perspectives. The test platform is a 6U CubeSat hosting an electric propulsion system and able to manage a variety of electric propulsion systems. The platform can regulate and distribute electric power (up to 60 W), exchange data according to several protocols (e.g., CAN bus, UART, I2C, SPI), and provide different mechanical layouts in 4U box completely dedicated to the propulsion system. Moreover, the data gathered by the onboard sensors are combined with the data from external devices and tools providing unprecedented information about the mutual behavior of a CubeSat platform and an electric propulsion system.

scholarly journals Laser–Accelerated Plasma–Propulsion System

2021 ◽  
Vol 11 (21) ◽  
pp. 10154
Author(s):  
Daniele Palla ◽  
Gabriele Cristoforetti
Keyword(s):  
Electric Propulsion ◽  
Laser Intensity ◽  
Laser Energy ◽  
Specific Impulse ◽  
Propulsion System ◽  
Relativistic Velocity ◽  
Plasma Propulsion ◽  
Propulsion Systems ◽  
Engine Thrust ◽  
Light Ions

In this paper, the laser-accelerated plasma–propulsion system (LAPPS) for a spacecraft is revisited. Starting from the general properties of relativistic propellants, the relations between specific impulse, engine thrust and rocket dynamics have been obtained. The specific impulse is defined in terms of the relativistic velocity of the propellant using the Walter’s parameterization, which is a suitable and general formalism for closed–cycle engines. Finally, the laser-driven acceleration of light ions via Target Normal Sheath Acceleration (TNSA) is discussed as a thruster. We find that LAPPS is capable of an impressive specific impulse Isp in the 105 s range for a laser intensity I0≃1021W/cm2. The limit of Isp≲104 s, which characterizes most of the other plasma-based space electric propulsion systems, can be obtained with a relatively low laser intensity of I0≳1019W/cm2. Finally, at fixed laser energy, the engine thrust can be larger by a factor 102 with respect to previous estimates, making the LAPPS potentially capable of thrust-power ratios in the N/MW range.

Flexible variable-specific-impulse electric propulsion systems for planetary missions

2006 ◽  
Vol 59 (8-11) ◽  
pp. 931-945 ◽  
Author(s):  
E. Chesta ◽  
D. Estublier ◽  
B. Fallis ◽  
E. Gengembre ◽  
J. Gonzalez del Amo ◽  
...  
Keyword(s):  
Electric Propulsion ◽  
Specific Impulse ◽  
Propulsion Systems ◽  
Planetary Missions
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