VASIMR Plasma Rocket Propulsion
Variable Specific Impulse Magnetoplasma Rocket (VASIMR)


VASIMR ELECTRICAL POWER REQUIRED FOR A 39 DAY TREK TO MARS
NEW! GO ...
BASIC ELECTRO-HYDRODYNAMIC THRUSTER (EHDT)
DESIGN GO ...
ION ENGINE DESIGN FOR TRAVEL IN OUTER SPACE GO ...

VASIMR Concept Trip to Mars VASIMR Concept Trip to Mars

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ELECTRICAL POWER REQUIRED FOR A 39 DAY TREK TO MARS
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Earth-Mars elliptical transfer orbitStarTravel has been used to perform a general heliocentric orbital transfer analyses to determine velocity increment (dV) and flight time (dT) required for traveling from Earth to Mars in 39 days as proposed in NASA's new budget. This combination MathCAD and StarTravel analysis uses an iterative solution based on the number (N) of 200 killowatt VASIMR plasma motors required to achieve the dV for Mars orbital insertion. A solution for traveling to Mars in 39 days is achieved when the equations of motion for gravity-free and drag-free space flight predict a total flight distance equal to the distance traveled along the elliptical transfer orbit predicted by StarTravel.

In summary, the following analysis predicts that a 12 megawatt power source is needed to energize a cluster of 60 VASIMR rocket motors to achieve the thrust required for sending a 37 metric ton spaceship to Mars in 39 days. This power source estimate is in line with previous statements by Dr. Franklin Chang Diaz who stated in a paper called The VASIMR Rocket that appeared in the November 2000 issue of Scientic American that a 10 to 12 megawatt nuclear reactor is required for a 39 day journey from Earth to Mars. In addition, on September 29, 2009 Dr. Franklin Chang Diaz stated the following. "In fact, with the power close to what a nuclear submarine generates, you could use VASIMR to fly humans to Mars in 39 days."

39 day transfer orbit to Mars
Figure-1, StarTravel results illustrating transfer orbit and orbital elements for a 39 day trip to Mars. Image size reduced.
 

MATHCAD ANALYSIS FOR DETERMINING VASIMR ELECTRICAL POWER,
THRUST AND MASS FLOW RATE FOR A 39 DAY TREK TO MARS
NEW!

A NEW VERSION OF StarTravel THAT INCORPORATES
THIS VASIMR PLASMA ROCKET ANALYSIS IS NOW AVAILABLE

Trip to Mars using the VASIMR engine

For validation of these results please visit the NASA VASIMR web site. Click Back to return.


BASIC ELECTRO-HYDRODYNAMIC THRUSTER (EHDT) DESIGN
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Ionization processThe ion engine design represented here is based on the concept of the Electro-Hydrodynamic Thruster (EHDT) which consists of a small diameter wire loop called the Sharp Edge Electrode (+V) separated from a three-sided sheet of aluminum foil called the Smooth Edge Electrode (-V). Where, either balsawood or plastic insulation material are used to separate the positive and negative electrodes. To start the process, Nitrogen and Oxygen gases in the vicinity of the positively charged wire loop are ionized when electrons from the supplied air are attracted to the positively charged electrode. The relatively small mass of positively charged ions are accelerated through the air gap between the two electrodes transferring linear momentum to the neutrally charged air. Neutrally charged air moves with the positively charged ions during the momentum transport process down the axis of the ion engine providing a very small thrust. Finally, positively charged ions are neutralized upon impact with the Smooth Edge Electrode and exit the Smooth Edge Electrode contributing to total ion engine thrust.

For the ion engine illustrated above the small diameter loop is replaced by a Sharp Edge Electrode disk having two very sharp internal and external circular edges for ionizing Nitrogen and Oxygen as air flows through the ion engine. In addition, the aluminum foil electrode is replaced by a very smooth aluminum cylinder having a uniformly rounded inlet edge that prevents coronal discharge and allows closer positive and negative electrode spacing. The illustration above displays the Smooth Edge Electrode as machined from a tube of aluminum in close proximity to the Sharp Edge Electrode. To power the ion engine a high voltage transformer from a Plasma Globe was used to supply 5kV to the positive Sharp Edge Electrode disk. A properly sized high voltage diode (8kV, 10 mA) was used to rectify voltage from 5kV AC supplied by the Power Globe transformer to 5 kV DC required by the positive electrode.

The potential advantage of the Smooth Edge Electrode is the elimination of a cathode neutralizer normally required to neutralize the positively charged plume of ions as the plume exits the ion engine by supplying a stream of electrons to the plume. Also, it may be the basic shape of these electrodes prevents the erosive effects normally observed on ion engine electrodes over long periods of time. This project is simply a demonstration for the use of the acceleration electrodes depicted in these illustrations.

Figure-1
Figure-1,
Plasma Globe powered ion engine and the mechanism
used to observe air flow through the ion engine device
Figure-2
Figure-2
Close up view of the high voltage power supply and high voltage diode used to rectify voltage from 5kV AC to 5kV DC
 
Figure-3
Figure-3
Close up view of the Sharp Edge Electrode (+5kV)
in front of the cylindrical Smooth Edge Electrode (-6V)
Figure-4
Figure-4
Back view of the ion engine showing the Smooth
Edge Electrode behind the Sharp Edge Electrode
 
Figure-5
Figure-5
Power to ion engine turned off - sheet stationary
Figure-6
Figure-6
Power to ion engine turned on - sheet deflected


ION ENGINE DESIGN FOR TRAVEL IN OUTER SPACE TOP
The high voltage transformer from the Plasma Globe pictured above is to weak to provide any meaningful ion engine thrust. In other words a standard design EHDT device powered by one of these Plasma Globe high voltage transformers could never levitate or even move around by itself on the end of a tether. To overcome these severe power limitations a second Plasma Globe high voltage transformer was added to excite a make-shift electron gun for an added source of electrons to ionize the fuel (Oxygen, Nitrogen and Argon) in the air. The addition of an electron gun worked rather well increasing sheet deflection and therefore ion engine thrust by a factor of approximately 5. Ion engine thrust was determined by applying the linear momentum equation for an inertial control volume of a free-jet impinging on the vertical sheet of paper. Ion engine thrust was also determined by measuring the amount of force required to deflect the vertical sheet of paper on the end of two threads as illustrated in Figure-1.

If this were an ion engine meant to operate in space the surrounding air will be replaced by Argon (Ar) injected at the front of the ion engine. Also, the ion engine design illustrated above will be enclosed by a circular cylinder so the ionized fuel (Argon) can circulate through and around the electrodes. Then, you would be off to the races, although rather slowly at first because ion engine thrust is on the order of only a few milliNewtons. A preliminary electrostatic thruster analysis indicates this design provides a specific impulse (Isp) of about 7,000 seconds. More work needs to be done on the preliminary ion engine design presented here to finalize the AeroRocket ion engine design. A final ion engine design is presented in Figure-12 that more closely resembles a design intended to operate in space. If this information proves useful or interesting please send an email to let us know.

Figure-7
Figure-7
Modified ion engine turned off - sheet stationary
Compare to sheet deflection in Figure-5
Figure-8
Figure-8
Modified ion engine turned on - sheet deflected
Compare to sheet deflection in Figure-6
 
Figure-9
Figure-9
Modified ion engine turned on - sheet deflected
Figure-10
Figure-10
System-view of the ion engine - sheet deflected
 
Figure-11
Figure-11, Top-view of the AeroRocket ion engine
 
PROPOSED ION ENGINE DESIGN
NASA plasma rocket

Using the electrostatic equations on page 590 of Rocket Propulsion Elements, 6th Edition the rocket motor thrust for this ion engine is 1.832 mN and the specific impulse is 19,000 seconds.

 
AeroRocket ion engine design
Figure-12, Front View, Back View and Side View of the AeroRocket ion engine
 
Ion engine thrust measurement system
Figure-13, Ion engine test system, turned off
Ion engine thrust measurement
Figure-14, Ion engine turned off (left) and on (right)

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| MAIN PAGE | PRODUCTS | CONSULTING | MISSION | RESUME |
Copyright 1999-2012 John Cipolla/AeroRocket. All rights reserved