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Electric Propulsion: Which One For My Spacecraft?
Ian J. E. Jordan Submitted December 6, 2000 to V. Pisacane as part of requirements for 744 Space Systems I course at JHU, Whiting School of Engineering.

1.0 Introduction
Electric propulsion has become a cost effective and sound engineering solution for many space applications. Two of the main reasons why are its increased commercial availability and the opportunity it affords to perform the same task as conventional chemical propulsion systems while reducing the portion of the spacecraft's mass required for that task. Electric propulsion systems have been tested on ground and in space since the 1960s, and a wide variety of system types are available or have flown. In this survey systems compatible with solar electric propulsion and that are commercially available are emphasized, however other types are mentioned for completeness. Blind pursuit of completeness can be a never ending endeavour and therefore some techniques will not be broached such as solar photon sailing, fusion, railguns, and laser propulsion. Included are summaries of the principles, advantages, disadvantages, ground based test experience, spaceflight experience, and descriptions of electric propulsion systems currently available. The reader should use this article as a pointer to other documentation sources as it is not meant to be an authoritative source of principles, design requirements, constraints, or data. Be also cautioned that research for the article tapped the resources of the world wide web extensively, and many of the references point to pages which have half-lives shorter than academic cycles. Occasionally, more conventional references are cited alongside URLs. With a control systems engineer in mind, an attempt is made to provide enough information for making preliminary choices about which type of electric propulsion might be suitable for an application. In this regard, the tables provide much of the "meat and potatoes" content. According to the Chemical Propulsion Information Agency, over 300 electric thrusters had flown on over 100 spacecraft as of 19971. In 1998, at least 78 more spacecraft used some type of electric propulsion device. By latest counts, 388 electric thrusters are aboard 152 spacecraft2. Electric propulsion research is an active field going as far back as the 1920s. Flight and test experience is mentioned briefly where the literature provides data.

1. Filliben, J. D., Johns Hopkins University, http://www.jhu.edu/~cpia/electhsys.html . 2. Snyder, J. S., Aerospace America, v. 38, no. 12, Dec., 2000.
Electric Propulsion: Which One For My Spacecraft? December 7, 2000 1


2.0 Classification of Electric Propulsion Types
With the wide variety of electric propulsion types, classification is in order to better understand them. Many primary authors have attempted to classify electric propulsion systems. It should be noted that not all authors agree on how each type of electric propulsion system is classified. For example, some authors classify Hall effect thrusters as electromagnetic while others classify it as electrostatic. The most common classification scheme is by the means in which the working fluid is accelerated up to exhaust velocity. However, alternate schemes divide the techniques by the means for ion production (for those that rely upon plasma generation) or by the propellant composition. Since different techniques can quite often each use various propellants, the latter scheme is not widely used. Although we discuss primarily solar electric propulsion, some systems work more efficiently at high power input and therefore operate better with nuclear or thermonuclear power sources. We adopt the conventional classification scheme wherein the primary category is how the working fluid is accelerated, while the secondary category is how the fluid (ionization in the case of plasmas) is generated. This scheme is fairly standard, and the three major categories of electric propulsion are: · · · Electrostatic Electromagnetic Electrothermal

Within these categories, subdivision by the techniques for ion generation (for those where plasma is generated) or by propellant heating is made. · Electrostatic · Ion Bombardment · Colloid ion · Contact ion · Field Emission (FEEP) · Microwave or Radiofrequency ion · Plasma separator ion · Radioisotopic ion · Hall Effect: · SPT · TAL · Electromagnetic · Magnetoplasmadynamic · Steady state vs Quasi steady state · Self-field vs Applied-field · Pulsed plasma · Helicon plasma · Inductive pulsed plasma · Electron-cyclotron-resonance · Variable specific-impulse plasma

· Electrothermal · Resistojet · Arcjet · DC vs AC vs Pulsed

· Electrothermal · Electrothermal hydrazine · Microwave electrothermal · Pulsed electrothermal

Table 1 shows the major techniques which have been researched, tested, and/or flown. Current developmental and historical information on the various types of electric propulsion systems is given.
Electric Propulsion: Which One For My Spacecraft? December 7, 2000 2


Table 1: General Features and Development Activity on Various Electric Propulsion Systems
Propulsion Type Flight Working Propellant Experience First Flight
7+ never never multiple never never H2, N2H4, N2, Teflon, dozens NH3, H20, N2O Xe, Ar, Ne, Cs, Hg Xe, Kr Glycerine Ce Ce Ce multiple 1+ never ? never never 7 ? 1970, SERT-II 1993 -

Commercial Level
Off-the-shelf Experimental Experimental Off-the-shelf Laboratory

Manufacturers or Test Labs
Centrospazio, IRS, Primex ? ? TRW, Primex, GRC Penn State, Michigan State, JPL, GRC Primex, Moog, EPL Hughes, GRC, JPL, AEA/ RAE, MMS, Japan DASA, Japan

Electrothermal DC Arcjet N2H4, H2, NH3, Teflon Electrothermal AC Arcjet Electrothermal Pulsed Arcjet Electrothermal Hydrazine Hydrazine Microwave Electrothermal N2 Pulsed Electrothermal Resistojet Ion Bombardment RF/Microwave Ion Colloid Ion Radioisotopic Plasma Separator Contact Ion FEEP Hall: SPT Hall: TAL

Developmental ? Off-the-shelf Off-the-shelf

1993 EURECA Experimental Flight Test 1962 1991 MIR 1972 1972 Theoretical Discontinued Discontinued Available Available Available

Developmental U.S., U.K. U.S. U.S. Hughes, LeRC, USAF Centrospazio, SRI, ESA, ARC Centrospazio, Fakel, AFRL, TsNIIMASH, NIITP, IST, RIAME GE, JPL, GRC, U. Tokyo, ISAS

liquid metal, Ce, Ru, In 5+ Xenon Xenon dozens dozens never

Electron-Cyclotron Reso- Ar, Kr, Xe nance

scheduled 2002 Laboratory / MUSES-C Verification EPEX on SFU 1985 1964, Zond 2 -

MPD Arcjet Steady State Nobles, H2, CH4, N2, never N2H4, NH3, Li, K, Na MPD Arcjet Quasi-Steady Nobles, H2, CH4, N2, ? State N2H4, NH3, Li, K, Na Pulsed Plasma Helicon Pulsed Plasma Pulsed Inductive N2H4, CO2, NH3, Ar Variable Specific Impulse Hydrogen Plasma MEMS Mini-Magnetospheric Plasma (M2P2) Teflon, Xe multiple never never never

Experimental / Centrospazio, EPPDL, Developmental IRS, MAI, ISAS Flight Tested Available Laboratory Laboratory ISAS EPPDL, Primex, NASA/ LeRC, SRL, Russia U. Wisconsin JSC, MIT

Developmental TRW

Pb-styphnate/Nitrocel- never ? lulose Solar Wind never

Laboratory / EPPDyL, AFRL Developmental Laboratory U. Washington

"Commercial Level" must of course be considered in the context that some systems may be at more
Electric Propulsion: Which One For My Spacecraft? December 7, 2000 3


advanced levels than indicated, and the list of providers is not intended to be complete. The preferred, tested, or commercially used propellants are shown, though quite often other options exist. Any noble gas can replace any other, although particular ones are preferred due to the need to match elemental properties with thruster performance. Cesium and mercury were tested and even flown in the 1960s in types of electrostatic systems that currently employ noble gases. Teflon can be replaced by other soft solids. Though not mentioned further here, C60 fullerenes are being considered as propellant, though significant problems remain with their use.

3.0 Electrostatic Propulsion
Electrostatic propulsion uses a high voltage electrostatic field to accelerate ions to large exhaust velocities. Most development work has been on systems using positively charged ions as the primary working fluid with some means for neutralizing the ions after they reach the exhaust. Many electrostatic systems rely on a gridded system at the exhaust port for containing and producing the high electric field needed to accelerate the ions (ion bombardment, RIT, and colloid thrusters). In a typical gridded thruster, a DC potential difference (~ 1 kV) between an inner grid anode on the plasma chamber and an exit-plane cathode is used. Multiple grids are employed at the exhaust port to divide the functions of propellant containment, ion acceleration, and beam divergence control3. Figure 1: Conceptual illustration for Hughes XIPS Radiative heat losses from the propellant heating/ ionization chamber limit efficiency in most systems. The power conditioning requirements are considered to be complex depending upon the method used for propellant ionization. Grid-type electrostatic systems have been lifetime limited by grid and cathode erosion. Material improvements, particularly with carbon-carbon composite grids are in development and expected to increase the performance and life expectancy of these units.
ion thruster showing major subsystem components.

3.1 Ion Bombardment
Electron bombardment or ion-bombardment thrusters produce positive ions by bombarding neutral propellant atoms in a discharge chamber with thermionically excited electrons. The discharge chamber is typically a cylindrical anode, with a centrally located axial hollow cathode. Heating of the axial

3. XIPS Figure 1 is from http://www.hughespace.com/factsheets/xips/nstar/ionengine.html .
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cathode produces thermionic emission of electrons at a low amperage (1 to 15 Amps) and low voltage (25 to 30 V DC) which discharge toward the outer anode. A magnetic field is typically applied in the discharge chamber to increase the electron path length and residence time allowing a higher collision probability and propellant utilization efficiency4. Grids held at different potentials near the exhaust port allow the plasma to be accelerated to exhaust velocities. The ions are then neutralized in the exhaust by a spray of electrons from a neutralizer cathode to prevent a potential difference from pulling ions Figure 2: Line diagram showing functional elements, fields, back to the engine. Ion bombardment systems have a test and flight history stretching over 30 years in space and 40 on the ground. Several NASA and Air Force low Earth-orbit missions in the 1970s helped demonstrate and troubleshoot the basic design. Intelsat VII employs 25 mN UK-10 northsouth station keeping units. Higher thrust units have been flight qualified for use as east-west stationkeeping motors on geosynchronous satellites (about 20 times the -V per year is required for eastwest versus north-south stationkeeping at geosynchronous altitudes). Perhaps the most prominent mission to date has been Deep Space-1 which uses a 30-cm diameter NASA/Hughes xenon-ion engine. Ground units have been tested at thrusts over 1 N, with power consumptions of 50 W - 200 kW. Many countries and organizations have been involved with development work and production of flight equipment including Glenn Research Center (GRC, formerly Lewis Research Center--LeRC), JPL, Hughes, and centers in Japan, German, France, UK, and Russia. Gridded acceleration can sometimes be tempermental as evinced with even the recent systems. DS1's engine had to initially be cycled hundreds of times before any thrust occurred to remove contaminants which were producing shorts in the acceleration electrodes. The initial flight of an electronbombardment test system (27mN SERT II mission in 1970) suffered similar problems. In addition to the drawbacks of gridded accelerators, the discharge chambers can be subject to sputtering and spalling. In spite of these problems, the relatively high performance, extended development heritage, and potential long lifetime of the thrusters make the systems attractive. Ion bombardment systems are capable of delivering the highest integrated lifetime impulse of any currently flying electric propulsion technique and are therefore excellent candidates for not only station keeping but for primary propulsion. Plans for commercial use are extensive. Tiny microthrusters which use a hollow cathode have been under test5.
and propellant paths of a gridded electrostatic ion bombardment thruster.

4. Electrostatic gridded element diagram Figure 2 from http://sec353.jpl.nasa.gov/apc/Electric/06.html 5. Katz, I., Davis, V. A., et. al., http://apc2000.jpl.nasa.gov/proceedings/Micr_PR1.pdf .
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3.2 Radiofrequency (RIT) & Microwave Ion
Radiofrequency ion thrusters rely on ion creation by pumping a cavity with radio frequency radiation, usually in an insulated discharge chamber. The ions are then extracted through the exhaust port by an accelerator grid similar to that in gridded electrostatic propulsion. Like other systems which eject a charged plasma, an external neutralizer cathode is used to balance the current in the exhaust. Thrust levels in the range of tens of milliNewtons have been demonstrated with the RIT-10 developed by the University of Giessen6 and Daimler-Benz Aerospace (DASA). The RIT-10 was flown aboard the EURECA test satellite in 1993. Larger models with 15-cm and 35-cm grids have been developed. ESA is pursuing investment in radiofrequency electroFigure 3: Conceptual diagram of the RIT thruster series showing static engines for use as north-south propellant flow and main thruster elements. station keeping for geostationary satellites. Two RIT-15s are being integrated for launch with the ARTEMIS satellite. RIT systems show lifetimes nearly equivalent to those achievable with the ion bombardment technique.

3.3 Electrostatic Colloid
With an electrostatic colloid thruster, droplets of a conductive liquid such as glycerol or sodium iodine are pumped through a needle at a high potential (~5-10 kV). An negatively charged extractor (several kV Figure 4: Gridded acceleration of colloid droplets is shown in negative potential) pulls the liquid into a the above layout sketch. thin continuous stream until positively charged droplets separate from the stream7.

6. Figure 3 showing the RIT design from http://www.irs.uni-stuttgart.de/RESEARCH/EL_PROP/ION/e_ion.html .
Electric Propulsion: Which One For My Spacecraft? December 7, 2000 6


Alternate propellant schemes rely upon condensing a supersaturated vapor, such as mercury or aluminum chlorides into a liquid. A drawback with the technique is that energy is required to condense the vapor, thereby reducing efficiency. Another is that droplet size is non-uniform resulting in inefficient drop mass/charge ratios. Studies the late oping a satellite in the 1960s and 1970s were performed in the U.K. and U.S., but development lagged until 1990s, largely because other systems appeared more promising. Busek Corporation is develcolloidal thruster system in the 25 microNewton class for the SBIR Air Force early warning system. Stanford University is also involved in testing small colloid thrusters8.

3.4 Contact Ion
The contact ion engine creates ions by passing liquid cesium through a heated bed of porous tungsten. The higher electron work function of tungsten relative to cesium results in electron transfer from the cesium atoms to the tungsten. The cesium ions vaporize from the tungsten bed and are accelerated by an electrostatic field. Structural and fabrication problems prevented the successful design of contact ion engines above the 20 kW power range. Excessive thermal management requirements, cesium toxicity and contamination of the engine, and the high system complexity made the technology unattractive. After about 7 orbital and sub-orbital U.S. flight tests spread across the 1960s and termination of a French program in the mid-1970s, pursuit of contact ion thruster development has apparently disappeared9.

3.5 Field Emission Ion
Field Emission Electric Propulsion (FEEP) relies on a strong electric field, typically 8 - 15 kV, to directly ionize the surface of a working fluid, typically a liquid metal. This technique differs principally from colloid ion thrusters in that individual ions are produced rather than droplets. Slit or pinhole diameters are typically about one micron in size10. The technique has an inherently high efficiency because no heat is lost in ionization since the fluid is directly accelerated after ionization and extraction
Figure 5: FEEP concept showing the ion flow, thruster elements, and application of electric potentials.

7. 8. 9. 10.

Line diagram of colloid thruster elements Figure 4 from http://www.busek.com/ . Mahoney, J., Perel, J., http://apc2000.jpl.nasa.gov/proceedings/Micr_PR2.pdf . A flight history can be found at http://sec353.jpl.nasa.gov/apc/Electric/08.html#UK . Figure 5 diagram of field emission ion concept from http://www.centrospazio.cpr.it/FEEPPrinciple.html .
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Electric Propulsion: Which One For My Spacecraft?


from the bulk material/fluid. One of the advantages from a spacecraft control point of view is the reproducibility of short pulse times and small impulse bits. Thrust units can be quite small and have comparatively high specific impulse. For space missions which require counteracting the effects of atmospheric drag and solar radiation pressure, FEEPs are well suited. Centrospazio11 and Austrian Research Center12 have produced flight units used on numerous missions for testing, and is marketing them for science-sensitive station keeping missions such as LISA and TPF.

3.6 Plasma Separator Ion
The plasma separator ion thruster uses high density cesium vapor which is pumped through an array of hollow cathodes having converging-diverging nozzles. Ions are created by discharge between the cathodes and annular anode ring at the end of the nozzles. After ionization, a conventional electrostatic acceleration system produces high exhaust velocities. Nonuniform plasma flow from the cathode array was noted in lab tests. Development work in US was discontinued in the 1960s with no space tests13.

3.7 Radioisotopic Ion
Unlike the name's implication, thrust is produced with charged colloidal particles. A thin layer of fuel containing beta-decaying radioisotopes is spread over a large emitting surface with a net positive charge produced by the decaying radioisotope fuel (typically conceived as Cesium 144). A large potential difference between surface and space (~500,000 - 1,500,000 volts) is produced as a result. Electrons emitted from this surface are collected by a shield, and the potential difference and current are used to operate a high-energy colloidal accelerator. In many respects, this was more of a technique for using radioisotopes to produce a large potential difference and drive an engine, and is thus a power source. No working model was produced and safety considerations have probably prevented the idea from being developed due to extreme radiation hazards and complexities.

3.8 Hall Current
R. G. Jahn14 has identified a number of different fundamental ways in which the Hall current or Hall effect can be used for electric propulsion. All Hall thrusters are gridless accelerators which use the
11. 12. 13. 14. Centrospazio: http://www.centrospazio.cpr.it/Centrospazio6FEEP.html . Austrian Research Center: http://www.arcs.ac.at/E/EM/ultra . What little information is available was found at http://sec353.jpl.nasa.gov/apc/Electric/08.html . Jahn, R. G., Physics of Electric Propulsion, McGraw-Hill, 1968, pp.219-227.
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Electric Propulsion: Which One For My Spacecraft?


body forces on charges in crossed electric and magnetic fields. The most common configurations in use have an axisymmetric cavity in which a radial magnetic field is generated by electromagnets with the field pointing to (or from) the surrounding coil/chamber body from (or to) an inner magnetic pole. Lower potential differences are required to reach specific impulses close to those of ion bombardment systems. Some of the exhaust energy is not produced by the electrostatic field and therefore some authors classify the system as electromagnetic rather than electrostatic. Hall thrusters research and development is Figure 6: Concept diagram for SPT-type hall device showing being pursued widely by in Europe, Japan, magnets (yellow), cathode (green), and plasma chamber (blue). the U.S., and Russia. Russia has been a longtime leader in the field of Hall thrusters. Typical usage is for orbit raising and station keeping maneuvers. Two types of Hall thrusters exist: · · Stationary Plasma Thruster (SPT) Thruster with Anode Layer (TAL)

3.8.1 Stationary Plasma Thruster In the stationary plasma thruster, or "closed-drift" a DC electric field is established along the axis of the device with the anode located at the non-exhaust end of the ionization chamber and with the cathode located externally or at the exhaust end of the chamber. Electrons emitted from the cathode are pulled into the chamber by the applied electric field and have their circulation lifetimes increased through a force produced by E x B drift15. As a result, they are available to ionize the propellant injected from the non-exhaust end of the chamber through collisions for longer times. The applied electrostatic field then accelerates the ions into the exhaust flow. Additional electrons from the cathode flow into the exhaust stream to neutralize the charge flow. Russian built Fakel SPTs have had over 10,000 hours of space operation16.

15. Figure 6 concept diagram from http://www.grc.nasa.gov/WWW/onboard/EPO/sld011.htm . Other good diagrams are available on the web at http://ncst-www.nrl.navy.mil/HomePage/EPDMHall.html, http://www.stanford.edu/group/pdl/ EP/Hallcut/HallCut.html and from http://www.busek.com/. 16. Choueiri, E., Aerospace America, v.37, no. 12, Dec. 1999, p.68.
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3.8.2 Thruster with Anode Layer (TAL) "Thruster with anode layer" or "thruster with external layer" devices are configured somewhat differently in that the anodes are located downstream. The ion production region is positioned more externally than in SPT units. Most are generally smaller and have a lower erosion rate. TAL type Hall thrusters typically have twice the thrust-to-power ratio of ion-bombardment electrostatic systems, yet require only 300 - 400 V electric field potentials. The USSR had been using both SPTs and TALs for decades with thrusters flying on Meteor weather, military surveillance, and Express telecommunications satellites. With the waning of the cold war, Russian Hall thrusters have even been used on U.S. military research projects, such as STEX, launched in October 1998. A successful marketing campaign in partnership with companies in the U.S. has been undertaken.

4.0 Electromagnetic Propulsion
A principle advantage of electromagnetic techniques over electrostatic ones is that plasmas with higher temperature and density (typically by several orders of magnitude) can be confined and directed. This results in higher exhaust velocities and therefore more efficient use of propellant mass. All electromagnetic techniques rely upon production of a plasma which is accelerated to exhaust velocity by interaction with electromagnetic fields within and around the plasma production chamber.

4.1 Magnetoplasmadynamic (MPD)
The MPD arcjet evolved from electrothermal arcjets and magnetogasdynamic technology and is sometimes referred to as a Lorentz Force Accelerator. The general configuration is an axisymmetric chamber with a cathode running the length of the ionization region which carries a large current (typically ~ kiloAmperes). The potential difference between electrodes ionizes inflowing neutral gas. Once ionized, the plasma is accelerated by both Joule heating and electrodynamic forces. The current carrying plasma interacts with a magnetic field resulting in a Lorentz acceleration which expels the plasma. The Lorentz force provides the dominant acceleration mechanism. There are two different Figure 7: Self-field MPD schematic showing ways the magnetic field is produced: electric and magnetic fields in addition to · · Self Field Applied Field
December 7, 2000

cathode, anode, and nozzle.

Electric Propulsion: Which One For My Spacecraft?

10


In both types, the cathode-anode is charged with a DC power supply. The high temperature cathode produces most of the current through thermionic emission. An azimuthal magnetic field is generated by the current flowing along the cathode bar, helping to initiate the J x B force on the plasma. With the self-field system17 acceleration of the plasma occurs as a result of the magnetic field produced by the high current flow along the cathode. With an applied field Figure 8: Applied-field MPD thruster showing technique18, an external solenoidal magnetic field is applied to enhance acceleration and plasma confinement. the additional solenoid (green) plus fields and For higher temperature and energy plasmas, applied azi- charge flow. muthal fields are preferred for plasma confinement. With the addition of the applied field, a lower plasma current is required to produce the same propulsive power. Depending upon the architecture, an MPD thruster can be operated in either a steady-state or pulsed mode with demonstrated durations as short as a millisecond. The design of magnetoplasmadynamic thrusters allow operation in one of two different modes: · · Quasi-Steady State Operation Steady State Operation

With pulsed (quasi-steady state) operation, a capacitive system discharges across the arc allowing for higher currents and higher magnetic fields for a given average power level. Arc erosion is more severe in the pulsed arcjets. Unfortunately for low-power applications, efficiency has been a function of power input. Cathode lifetime is a limitation as with many systems employing high velocity plasmas. No steady state units have flown as these generally require higher input power. High thrust and specific impulse are achievable. The Japanese ISAS satellite SFU tested the general operation of quasi-steady state MPD thrusters.

4.2 Pulsed Plasma (PPT)
In a pulsed plasma thruster19, plasma is created by an arc discharge from a capacitor across a pair of electrodes. The ions in the plasma are then accelerated by the J x B Lorentz force in the induced magnetic field. Their use stretches back to the mid-1960s, and a resurgence in popularity has been seen in the 1990s. In a teflon pulsed plasma thruster, ablation of a solid block of teflon propellant. In these thrusters, propellant delivery is by quite a simple device: a spring! Teflon PPTs are currently
17. Self-field concept diagram (Fig. 7) from http://www.irs.uni-stuttgart.de/RESEARCH/EL_PROP/MPD/e_mpd.html . 18. Applied field concept diagram (Fig. 8) from http://www.irs.uni-stuttgart.de/RESEARCH/EL_PROP/AFMPD/e_afmpd.html . 19. Figure 9 teflon PPT cross-section diagram from http://www.lerc.nasa.gov/WWW/RT1997/6000/6910curran.htm .
Electric Propulsion: Which One For My Spacecraft? December 7, 2000 11


being used on EO-1 for momentum wheel replacement20 in pitch-axis control. Teflon PPTs are also a leading candidate for DS-3, an experimental space interferometer which requires precision formation flying. Use as orbit-raising, attitude control, station keeping, and fine pointing and positioning control will no doubt continue. Another type of pulsed plasma system under development uses a gas feed system with xenon or water as a propellant. This system has not yet flown but has been shown to be capable of over 2 million cycles at a power consumption rate of 40 W and 0.02 second pulses using water as a propellant.

Figure 9: Teflon pulsed plasma thruster cross-sectional diagram. Relative locations of the capacitance, propellant, ignition, and electrode positions are shown.

4.3 Helicon Plasma
A Helicon thruster21 is similar to pulsed plasma thrusters in that acceleration of a plasma occurs through a Lorentz force interaction. However, a helicon tube is instead used to produce a travelling electromagnetic wave down the center of the plasma chamber to maintain the high magnetic field strengths. Helicons have also been proposed for use in other electromagnetic thrusters, such as in the variable specific impulse plasma thruster.

4.4 Inductive Pulsed Plasma (PIT)
In a pulsed inductive thruster22, a bank of high voltage (~ 10 kV) capacitors is discharged into a flat induction coil strip surrounding a cylindrical chamber. Immediately before discharge, a fast valving system injects neutral propellant into the chamber. Before the propellant can escape the chamber by diffusion, the discharge pulse creates a high magnetic field within the chamber which both Figure 10: PIT conceptual diagram showing the inductive curionizes and accelerates the plasma away from rent strip (green), the insulating chamber (orange) and the
plasma (red).

20. Burton, R. L., Aerospace America, v.36, no. 12, Dec. 1998, p.62. 21. Source info for the Helicon Plasma Thruster is available at http://rigel.neep.wisc.edu/~jfs/neep533.lect31.99/plasmaProp.html , and information about helicons at http://www.anutech.com.au/asi/helicon.htm . 22. Figure 10 cut-away diagram from http://sec353.jpl.nasa.gov/apc/Electric/16.html .
Electric Propulsion: Which One For My Spacecraft? December 7, 2000 12


the coil. Since the chamber is closed on one end, the propellant plasma is confined and squeeze out of the exhaust end of the chamber. Shielding transient magnetic and electric fields may be an issue. Development work by TRW has been performed, but no space tests have been conducted.

4.5 Variable Specific Impulse Plasma
The technique relies on an electrodeless cylindrical chamber in which the functions of ionization, plasma heating, and conversion to a directed exhaust. In the forward portion of the chamber, hydrogen is injected and ionized. The ions diffuse into the mid-section of the chamber and are further heated by electron and ion cyclotron heating and whistler wave heating. Moving downstream, the plasma enters the nozzle section where the shape and field strength and configuration converts the plasma's thermal energy into propulsive kinetic Figure 11: Diagram showing relationship of the VASIMR propulsion and control elements. energy. The chamber walls are kept cool by injecting neutral hydrogen radially near the nozzle, with the bonus of disrupting the connection between the plasma and the contained magnetic field at low specific impulse (high plasma densities). At higher specific impulses, a high frequency AC ripple field is instead used to disrupt the magnetic field - plasma connection. The variable specific impulse plasma thruster is currently only in the laboratory testing and development phases. Johnson Space Center is currently exploring a version suitable for manned space missions called VASIMR23.

4.6 Electron-Cyclotron-Resonance (ECR)
ECR is an electrodeless technique using a m