NASA/TM--1999-209436

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AIAA 99-2949

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High Power Hall Thrusters

Robert Jankovsky

Glenn Research Center, Cleveland, Ohio

Sergey Tverdokhlebov

TsNIIMASH Export, Korolev, Moscow, Russia

David Manzella

Dynacs Engineering Company, Inc., Brook Park, Ohio

December 1999

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NASA/TM--1999-209436 AIAA 99-2949

High Power Hall Thrusters

Robert Jankovsky

Glenn Research Center, Cleveland, Ohio

Sergey Tverdokhlebov

TsNIIMASH Export, Korolev, Moscow, Russia

David Manzella

Dynacs Engineering Company, Inc., Brook Park, Ohio

Prepared for the

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cosponsored by AIAA, ASME, SAE, and ASEE

Los Angeles, California, June 20-24, 1999

National Aeronautics and

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Glenn Research Center

December 1999
NASA Center for Aerospace Information

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High Power Hall Thrusters

Robert Jankovsky

National Aeronautics and Space Administration

Glenn Research Center

21000 Brookpark Rd

Cleveland, Ohio 44135

Sergey Tverdokhlebov

TsNIIMASH Export

4 Pionerskaya St

Korolev, Moscow Region

141070, Russia

David Manzella

Dynacs Engineering Company Inc.

2001 Aerospace Parkway

Brook Park, Ohio 44142

Abstract

The development of Hall thrusters with powers ranging from tens of kilowatts to in excess of one hundred kilowatts

is considered based on renewed interest in high powel; high thrust electric propulsion applications. An approach

to develop such thrusters based on previous experience is discussed. It is shown that the previous experimental

data taken with thrusters of 10 kW input power and less can be used. Potential mass savings due to the design of

high power Hall thrusters are discussed. Both xenon and alternate thruster propellant are considered, as are

technological issues that will challenge the design of high power Hall thrusters. Finally, the intplications of such a

development effort with regard to ground testbtg and spacecraft integration issues are discussed.

Introduction

It has long been the goal of both US and

Russian scientists to use high power electric propulsion for primary spacecraft propulsion. Missions

to Mars, reusable space tugs, and other propulsion

intensive missions have all been considered since

the early 1960's. The opportunities to conduct these

missions have never arisen due to the lack of onboard spacecraft power. Additionally, due to a

combination of technical problems, political

pressure, and unfortunate timing the possibility for

space nuclear power has been virtually eliminated.

As a result the major electric propulsion

development efforts over the last several decades

has concentrated on devices with powers of only a

few kilowatts or less.

With regard to Hall thrusters, these efforts,

which were for the most part conducted in what

was then the Soviet Union, have resulted in

successful operational deployment of these devices

for stationkeeping purposes on Soviet and later

Russian spacecraft. The early engines operated at

0.6 kW and later engines operated at 1.35 kW.

Since the end of the cold war this very successful

Hall thruster technology became available to the

rest of the world. The first implementation of this

technology on a Western spacecraft occurred in

1998 when a device of the anode layer type

(TALk derated for operation at 0.6 kW, was flown

on a U.S. government experimental spacecraft.

There are a multitude of additional spacecraft

being planned that will use higher power Hall

thrusters in the near future. The highest power

system being developed for near term flight is a

5 kW system utilizing a TI60E SPT type engine

slated for flight in early 2000.

Spacecraft power system designs have steadily

evolved. Spacecraft with power levels approaching

20 kW are currently being implemented. As a

result, electric propulsion, and specifically Hall

thrusters, are being considered once again for high

total impulse missions that have not been

considered since the possibility for space nuclear

power died.

The combination of advanced power systems

and advanced orbital trajectories have mission

designers emphasizing a significant need for high

power Hall thruster technology with a high ratio of

thrust-to-power. This technology promises the

possibility of significant reductions in launch mass

making new missions possible with the existing

launch vehicle fleet. Non-nuclear trans-lunar and

NASA/TM--1999-209436 1
trans-Marsinjections,spacesolarpowersatellites,

SpaceStationReboost.LEO-GEO transfers in less

than 90 days are a few examples of these new

missions._-"34 The thruster power levels required for

these missions range from 30 to 100 kW, in

excess of what is currently available. As a result,

the objective of this paper is to give preliminary

consideration to the general issues associated with

the development of high power Hall thrusters, suggest reasonable limits for near term extrapolation

of state-of-the-art (SOA) Hall technology, and to

suggest a development path for the even higher

power thrusters.

Nomenclature

T thrust

P power

I_ ion current

J_ ion current density

I,t,, permeability of free space

_ average ion exhaust velocity

m mass flowrate (use typical sign with upper dot)

e electron charge

M, molecular wt. ion

B magnetic field strength

IH Hall current caused by electric drift of

electrons

D average diameter of thruster channel

r_ ion Larmour radius

V_ discharge voltage

b,. channel width

Ij discharge current

j_. electron current density

k_ constant

k, constant

Sch channel cross section area

I m equivalent mass flow rate in current units

I_,.. electron cyclotron frequency

ag_.,,electron-neutral collision frequency

Background

There is limited published data relating to the

performance of high-power Hall devices. Those

data that are available were taken with laboratory

model hardware and are generally not comprehensive in nature. However, a review of this work

indicates the current state-of-the-art with respect to

high power Hall thrusters. Beginning in the early

1980"s laboratory models of high power xenon

stationary plasma thrusters with outer discharge

chamber diameters up to 290 mm were developed

and tested in Russia at Fakel in cooperation with

Moscow Aviation Institute (MAI) and the

Kurchatov Institute. __ At the same time large-scale

anode layer thrusters (TALs) were developed and

tested in Russia at TsNIIMASH in cooperation

with RSC "Energia." Initially these TAL tests

investigated bismuth as a propellantfl Later tests

considered xenon as the propellant. In both cases

the objective of these tests was to demonstrate

operation with high specific impulse/ More

recently the performance of one of these high

power TALs developed at TsNIIMASH for operation on xenon, given the designation TM-50, was

evaluated at NASA GRC for operation in a high

thrust, low specific impulse modefl Most recently,

Jankovsky et al. reported the preliminary results

from testing an SPT type thruster, designated the

T-220 at NASA GRC. _' This thruster was

developed in the United States by TRW in

cooperation with Space Power Incorporated. A

portion of these data encompassing the range of

operation for each these devices is summarized in

the Table 1 and shown graphically in Figure 1.

Physical Limitations

The operational features of low-power thrusters

with closed electron drift have previously been

comprehensively investigated. Therefore the

analysis below focuses on the physical constraints

limiting the maximum power of a single Hall

thruster. Arrays of lower power Hall thruster are an

alternate strategy for obtaining high power Hall

thruster propulsion systems, however neither this

approach nor multi-channel Hall thrusters are

considered in this analysis. Both of these

approaches may offer certain advantages, but they

are evident design derivatives from the single

thruster channel which is considered.

The general desire to evaluate high-power Hall

thrusters assumes their capability to operate with

the highest possible thrust-to-power ratio. Because

this ratio is inversely proportional to the average

ion exhaust velocity, T/P~I/'o i, operation with

relatively low specific impulse, Isp, at low

discharge voltages is preferable. Thus a thruster's

ability to accelerate a maximum ion current I i is of

great interest. Correspondingly for a thruster of

given cross section, the maximum achievable ion

current density Ji represents a physical constraint

which must be considered.

The maximum permissible value for J, in Hall

thrusters can be estimated by utilizing the

magnetohydrodynamic approximation for the

momentum flux density of ions accelerated by Hall

current. This approximation, first suggested by A.

Zharinov and Y. Popov over thirty years ago,

neglects electron pressure and the influence of

self-induced magnetic field:

NASA/TM--1999-209436 2
JiM_xh/e<_ B2/2JJo

Therefore, to increase ion current it is

necessary to increase the cross section of the

thruster channel or increase the magnetic field

strength. When the magnetic field is increased to

satisfy this condition the ion current density is

proportional to the Hall drift current In:

Ji = Ire/2 l-d)ri

Where D is the average diameter of thruster

discharge chamber and r, is the ion Larmour or

gyroradius.

Because of the electron losses to the walls.

anomalous electron mobility caused by the

scattering attributed to oscillations, and due to

azimuthal non-uniformity's, the Hall current value

is always less than its theoretical maximum.

X,,

I.= fj_6J,(1/o,.)a.,

0

Another potential limit to scaling of Hall

thrusters is the distortion of the external magnetic

field, caused by the diamagnetic influence of the

Hall current. If the self-induced magnetic field of

the Hall current is significant, under certain

conditions, a significant distortion of the electric

field distribution along the thruster channel and

subsequent unpredictable thruster performance may

develop. Although all aspects of this phenomena

are not well understood, the strong gradient of

magnetic field in the thruster channel may help to

minimize this distortion.

Another potential limit to the maximum

discharge current is Joule heating of the anode

caused by the back-streaming of electrons. The

average electron energy at low mass flow rates for

a TAL may be as high as eva/5. This requires up

to 20% of the input power be radiated or conducted

away from the anode. However, generally as the

neutral density increases the average electron

energy decreases, but the total power increases

and the anode energy dissipation remains constant

over a relatively wide range of thruster operation.

This is referred to as the "power plateau". Further

increases in propellant flow rate above this "power

plateau" results in a substantial increase in

electron back-streaming evident by large increases

in discharge current with incremental flow rate

increases. Substantial flow rate, or current

increase, in this regime of operation are generally

limited by the thermal design of the anode which

serves as a practical limit for high power operation

of a given device.

The choice of propellant also impacts the

necessary ion current density for efficient thruster

operation. While the previously presented equation

shows that a lower magnetic field is all that is

required to accelerate the same flux of lighter ions

as compared to a heavier ion, an acceptable value

of propellant utilization (or efficiency) requires

higher current density for the lighter propellantJ j

Thus, while the thermal limitations of a given

thruster are not affected by the choice of

propellant, the necessity to increase the discharge

current to maintain the efficiency while operating

with lighter propellants does reduce the margin

with respect to the thermal design. The increase in

discharge current also results in a reduction in the

attainable discharge voltage at fixed power.

Because these considerations are general in

nature they are applicable to both SPT and TALtype Hall thrusters. A quasi-neutral mode of

operation is assumed, and the so called vacuum

mode of TAL operation which has been previously

described by others _' is not considered primarily

due to the unfavorable thrust-to-power ratio

encountered during operation in this mode.

Design Parameters

Simple correlations between power and size of

large-scale Hall thrusters can be easily made

based on existing laboratory, engineering and flight

models of SPTs and TALs. To assess the target

performances for high power Hall thruster

development comparative analysis was made with

use of published data v-12131_15on power, characteristic size and mass of 6 SPT models and 5 TAL

models. For comparison convenience the values of

power and mass flow rates were taken from data

corresponding to an l_p of approximately 2500 sec.

Also, for convenience an average diameter of the

thruster channel was used instead of the outer

diameter which is the convention in SPT-related

publications. Thrusters with average diameter of

38-250 mm were considered.

Figure 2 illustrates the general scaling trend of

Hall devices as power increases with channel size.

For a discharge chamber with an average diameter

of 400 mm power levels of 37 kW and 48 kW

would be predicted for operation at 2500 second

specific impulse with an SPT and TAL

respectively. This result is consistent with the fact

that TAL is characteristically smaller device than

SPT for a given power level (for example an SPT100 and TAL D-55 have the same nominal power

capability). As is also seen in this figure with data

from TAL testing, by increasing the specific

NASA/TM--1999-209436 3
impulserequirementfora giventhrustersizethe

maximumpredictedpowerlevel also increases,

howeverattheexpenseof thrust.Thisdependence

of specific impulseon dischargevoltage is

predictablebasedonthemassflowrate.thruster

dimensions,and the exchangeparameter_,

definedas the ratio of the dischargecurrent

dividedby thecurrentequivalentmassflowrate

forasinglyionizedpropellant.

With somemathematicalmanipulationone

can write the thrusterpowerin termsof the

dischargevoltage,the exchangeparameter,the

thrusterdiameter,andseveralvariablesdepending

ontheareaandwidthof thedischargechamber:

P = kiD2(l+k2/D) _, Va .

where k_ _ reSen and k2 _ m/D

In SPT thrusters the parameter of exchange _,

has been shown to increase linearly with the

increased channel average diameter D and width

be. For instance, with an increase of average

diameter from 50 mm to 250 mm, the increase in

the exchange parameter from 0.9 to 1.4 was

reported. _6 However, the ratio of m/Sch remained

constant. Previously it has been suggested that the

ratio of tolD remains constant with thruster

size, _6_7 however, based on the data shown in

Figure 3 there is some variation with thruster size

and. therefore, it is believed that in general m/D¢

constant. More specifically, those cases in which

m/D is held constant are not of significant interest

because they require an increase of the thruster

diameter while holding other discharge chamber

parameters constant, thus sacrificing performance.

Finally, while it appears that for the most part

m/sch is constant it is possible to increase this

ratio at the expense ot" thruster lifetime.

Similar considerations may be used to

estimate thruster mass. The required magnetic field

strength B is directly proportional to the discharge

current la for a wide range of mass flow rates (for

discharge voltages in the range of 200-600 V). The

mass of the magnetic system usually represents

about 75-80% of total thruster mass. The difference

between the mass value for existing engineering

models and the targets for flight models was also

considered to estimate the thruster mass reduction

trend during the development efforts from laboratory to flight status. The values of thruster mass

were corrected to exclude cathode(s) and orifice

block. The results are presented in the Figure 4.

The mass reduction trend is generally similar

for both SPT and TAL, so the approximate values

of 50 and 40 kg may be predicted for both an SPT

and TAL 50 kW, 2500 sec laboratory and flightweight thruster. Thruster specific mass goes down

with the increase of power. Therefore, for

comparison purposes, it can be seen that a single

50 kW thruster would be 30'7c lighter than ten

5 kW thrusters.

A simple parametrical approach like the one

described above does not involve a thorough consideration of all the physical constraints such as

critical ion current density or thruster thermal

mode. Nevertheless, it is useful for initial engineering estimates and it gives an opportunity to

imagine the sequence of attempts to employ the

traditional Hall thruster design in the extreme

conditions. For instance, it may be shown that a

100 kW Hall thruster with a specific impulse of

1600 s (300 V for xenon) may be as large as one

meter in average diameter.

Technological Issues of Thruster Design

The development of high-power Hall thrusters

as well as of any large-scale device represents

certain design and technological challenges. While

there are numerous possible design configurations

for Hall-current devices only the most general

characteristics will be considered. First. a linear

increase in Hall thruster dimensions involves

certain difficulties with providing an azimuthally

homogeneous magnetic field and gas distribution

in the thruster channel required for normal drift

motion of electrons. Local non-uniformity's may

result in potential eccentricities in the thruster

plume. Potential reasons include, but are not

limited to, local changes in the discharge chamber

walls conductivity (issue of ceramic material

homogeneity), mechanical alignment, and cathode

location. These may lead to several effects such as

non-symmetric wear of the discharge chamber

walls and shifts in the thrust vector. All these

features that are generally inherent to any size of

Hall thrusters may be strongly amplified while

occurring in a large design.

Increase of thruster size may also complicate

thruster design from a thermal-mechanical standpoint. This is because absolute values of linear

expansions expected for thermal stressed elements

of the thruster are proportional to its average

diameter and are comparable with the characteristic length of acceleration zone. Another significant issue is the temperature gradients originating

across the channel walls. This will necessitate a

transient as well as steady-state thermal analysis

of critical components such as thruster ceramics. A

preliminary consideration has already been given

to transient analysis of high power hall thrusters. It

has been demonstrated experimentally that it may

take multiple hours to reach thermal equilibrium.

In one experiment a non-monotonic increase in

temperature was observed during a transient mode.

NASA/TM-- 1999-209436 4
Thisindicatespotentiallydifferentcontributionin

time of the thermalconductivityand radiation

mechanismsof heat exchange.Corresponding

changesin performancesmayaccompanythese

differentthermalregimes,whichtherefore,will

needto beaconcernof futurehighpowerdesigns.

Generally,if theapproximateheatflux distributionin the thrusterchamberis knownthe

simulationof athrusterthermalmodemaybedone

with useof finite elementsmethods.The only

problemis to determinea boundaryconditionson

heat-exchangingand radiatingsurfaces.As for

detailedthermalanalysisof a thruster,oneof the

remaininguncertaintiesis associatedwith the

correctdeterminationof theenergyassociatedwith

plumeemissionintheVUVrange.

Anotherissueis thedevelopment of a special

propellantinsulator.Becauseof increasedvoltages,largemassflowrates,highanodetemperature,presenceof highfrequencyelectricfieldand

residualmagneticfield,etc.,a newreliabledesign

is required.So,it is clearthattherearepractical

imitationswithregardhowlargehighpowerHall

thrusterscanbe made,basedon theindividual

componentsusedto makethesethrusterswhich

will haveto beaddressedin orderto permitthe

development ofhighpowerHallthrusters.

Propellant

Propellant selection for High Power Hall is of

critical importance. Xenon is the only propellant

under utilization in the current and near term

missions employing Hall thruster propulsion. While

typical magnitudes of propellant mass required for

projected telecommunication satellites with 1 to

5 kW Hall thrusters does not exceed 100 kg,

hundred-kW space missions would require tons of

propellant. Assuming total thrust time for two round

trips to a high elliptical point of trans-planet

injection, utilizing 500 kW of solar electric power,

a propulsion system consisting of 10 Hall thrusters

consuming 50 kW power each could be imagined.

Such thrusters might have a flow rate of 100 mg/s

of xenon each with a specific impulse of 2500 sec.

For this mission, a propellant mass of ~30 metric

tons is required. Availability of xenon supply,

therefore, becomes one of the major concerns.

Potential alternatives under discussion include

krypton, xenon-krypton mixtures or condensable

metal propellants.

From a purely thruster physics standpoint, it is

clear that to maximize thrust-to-power the best

choice of propellant is the highest molecular

weight element that that does not pose an

excessively high cost for ionization. A few

examples of alternate propellants are in Table 2.

For comparison purposes changes in thruster

efficiency with alternate propellants is not

included. While only considering the thruster

physics in selecting a propellant is not a

recommended approach, it is the basis from which

initial considerations should start. From the initial

thruster physics basis other considerations such as

propellant management, spacecraft contamination,

ground testing, and environmental impact can be

traded against trip time, spacecraft power system

and launch vehicle for each application. The trade"

for a geostationary spacecraft that desires to do

low Earth orbit to geostationary orbit transfer with

a Hall thruster or the International Space Station

which desires to use a Hall thruster for reboost may

be quite different than a manned mission to Mars

or a trans-Neptune stage.

If a propellant other than xenon is pursued

several issues will need to be addresses. Here we

will qualitatively consider only some of them for

the most commonly considered alternative,

krypton. The first is with respect to the impact of

utilizing krypton as a propellant which may seem

attractive based on it similarity to xenon. In

practice, however, the higher ionization potential

and lower atomic weight of krypton lead to

substantial decreases in Hall thruster performance.

Preliminary investigations conducted with singlestage SPTs and TALs operated with krypton

demonstrated 15-20% less efficiency than with

xenon, ts Even with a greater than 50% increase in

mass flow rate and subsequent increase in the discharge current the expected efficiency on krypton

in the specific impulse range of 2000-2500 sec will

probably not be in excess of 40-45%. The penalty

associated with the operation at the increased

discharge currents is a higher thruster thermal load

as compared with operation on xenon.

Mixtures of gases such as xenon and krypton

have also been previously considered. For our consideration, however, the extra energy dissipation in

the thruster structure may still be the limiting

technical issue because it constrains the maximum

input power of a thruster with given size. Certainly,

the addition of 10-50% molar impurity of xenon to

a krypton propellant will result in improvement of

the thruster performance as compared to pure

krypton. However, no principal "resonant"

phenomenon facilitating the increase of partial

krypton utilization efficiency with xenon are

expected. This is verified by limited experimentation which demonstrated the additive character of

xenon's contribution to thrust and efficiency.

Therefore. the attractiveness of use of the

NASA/TM--1999-209436 5
xenon-kryptonmixture"to resolvea cost and

propellantavailabilityissuesis still in direct

proportionof xenonfractionused.Ultimatelyeither

costorothersystemsconsiderationsarelikely to

drivethe propellantselection.However.at this

time preliminaryanalysisbasedon combined

considerationof cost/performances/system integrationimpactsshowsthat the mostreasonable

and cost effectiveapproachfor the nearest

developmentof high-powerHall thrusteris to

completethe evaluationand test programof

hardwareon xenon.Also a dedicatedanalysis

whichgivescarefulconsiderationto the benefits

vs.environmental impactsof development of metal

propellantsystemsshouldbeundertaken.

Test & Integration

The testing and integration of high power Hall

thrusters will have an impact on the results of any

design consideration for this type of thruster. Of

course, the magnitude of the impact is dependent

on such things as thruster type and propellant

selection (i.e.. non condensable versus condensable). Also. the possibility for things like plume

impingement onto spacecraft surfaces, or contamination of sensitive optical surfaces are highly

dependent on a particular geometry and spacecraft

configuration. In the following discussions these

issues will be considered in general for a single

thruster. If an array of thrusters is utilized

additional complications, beyond those discussed

here may be manifest.

Although it may be possible to design a Hall

thruster for a given application with favorable

operational characteristics in space, if the design

cannot be qualified and acceptance tested in an

acceptable fashion on the ground the design is of

questionable utility. As a result, the issues related

to the ground testing of large Hall thrusters are of

primary importance. The main technical consideration with regard to ground testing is what level of

vacuum is needed within the ground test facility to

adequately simulate the environment in which the

thruster would operate in space. The ability of a

test facility to sustain a given vacuum pressure is

related to its pumping speed. Typically a surface

which will condense the propellant upon it is used

as the pump. this often times requires cryogenic

temperatures for gases such as krypton and xenon.

Note that obtaining of a pure Xe-Kr mixture may be even more

expensive procedure than direct production ojXe fi'om initial

pro&tct of industrial air -separating facilities, containing

0.15..0.25 % qfxemm and ko'pton.

Previously, for consideration of propellant

selection, a xenon flow rate of 100 mg/s was considered as a possibility for a single 50kW thruster

with a specific impulse of 2500 sec. For thrusters

of the SPT type the facility pressure was shown to

have an effect on the measured performance at

pressures above 2x 10-_ torr. j5 In order for this effect.

thought to be caused by ingestion of background

gas into the discharge chamber of the engine, to be

insignificant, a xenon pumping speed in excess of

600,000 liters per second would be required. While

this pumping speed is within the range of currently

existing electric propulsion test facilities, those

with a order of magnitude higher pumping speeds

are not. Additionally, thrusters of the traditional

anode layer type tend to have a smaller cross

sectional area at the exit plane than do SPT type

thrusters. As a result it is likely that chamber

pressure does not have an effect on performance

until slightly higher pressures are reached.

These considerations have focused on the

effect of mass flow rate, which to first order determines the current in a Hall thruster and on vacuum

facility background pressure during testing. When

discussing high power operation the effect of

discharge voltage should also be considered. As

discharge voltage is increased the effect of the

ingested background gas on performance changes.

This is because at higher voltage, the ingested

neutral gas has a lower probability of diffusing

completely through the acceleration zone within

the discharge chamber before being ionized by

electron impact and subsequently accelerated

itself. Therefore these ingested, ionized, and subsequently accelerated background atoms do not

produce as much thrust as those propellant atoms

introduced at the rear of the discharge chamber

through the anode. This is the basis for the varying

effect of background pressure on performance with

discharge voltage. The over riding consideration is

that in any ground test the effect of the finite

background pressure must be adequately understood. Corrections can be made to account for this

effect, but if the corrections are not adequate a

real possibility exists for thruster operation in

space to be significantly different than during

qualification and ground testing.

A second consideration with regard to testing

of high power Hall thrusters is the ability to

adequately address relevant integration issues.

While these individual issues will subsequently be

considered, a significant number of these are

related to the distribution of effluent in the plume.

Background gas within a test facility during

thruster operation may charge exchange with

energetic plume ions affecting the distribution and

composition of the plasma generated by the

engine. While estimates of this effect can be made

NASA/TM--1999-209436 6
basedonknownchargeexchangecrosssections,

_brexample,experimental evidenceindicatesthat

evenatpressuresbelow2x10-_torrthisphenomena

isstillsignificantforthoseregionsof theplumeat

largeangleto the directionof thrust._ For the

same100mg/sHallthrusterconsideredabove,this

correspondsto a 6,000,000l/s requiredpumping

speedwhich,aspreviouslystated,is notcurrently

possiblein existing electric propulsiontest

facilities.

Anothertestingissueof seriousconsequence is

relatedto sputteringof thesurfacesubjectedto

direction impingementwithin the test facility.

Conceptuallythisissueis easyto visualize.The

engineisdesignedtoproducehighenergyparticles

whichcantravelat velocitieswell in excessof

20km/s.Dueto theextremeenergeticnatureof

theseparticles,uponimpactwithasurfacethereis

afiniteprobabilitythattheywill removeorsputter

a portionof theparentmaterial.Afterlongterm

bombardment of this type the magnitudeof

materialliberatedcanbecomequite substantial.

For this reason, sputter resistant materials such as

graphite are often used as beam targets. The

erosion of the facility itself posses no real consequence to the test as long as the actual

structural integrity of the facility is not compromised, however, the sputtered material will

redeposit elsewhere within the facility, including

on the test hardware with potentially adverse

results. This effect can be minimized through

careful design of targets and baffles, but is

certainly a significant issue with respect to the

long term testing of high power Hall thrusters. To

illustrate the magnitude of this effect, again consider the 50 kW thruster with a xenon mass flow

rate of 100 mg/s. For a graphite target at 5 m from

the exit plane of the engine the surface would be

eroded 0.1 mm after 5000 hrs of operation. If the

surface was stainless steel it would be eroded to a

depth of 0.6 mm. While this may not seem

substantial the total amount of graphite sputtered

throughout the tank would be as much as 46 kg and

for steel in excess of 1000 kg. Clearly the

redeposition of this material may be a significant

concern for the test hardware.

The integration issues associated with high

power Hall thrusters are not significantly different

than those associated with lower power Hall

thrusters, although the impact on a particular

spacecraft may be considerably greater due to the

larger size and higher fluxes. Direct impingement

of particles accelerated out the engine transfer

both momentum and thermal energy. For high

power thrusters this issue will likely be minimized

through spacecraft design, but this certainly will be

an integration issue worthy of consideration.

Another issue of potentially significant impact

from a spacecraft integration issue is the

accommodation of the waste heat generated during

thruster operation. A thruster with a reasonably

efficient thermal design may still need to reject

25% of its input power as thermal energy.

Minimizing the effect of the conducted and

radiated heat on the spacecraft will be a significant design challenge for the high power Hall

thruster spacecraft integration specialist.

There are a number of additional issues for

consideration from a spacecraft integration perspective. These include the possibility of an off

axis thrust vector which could vary in time. If an

engine does produce a time varying force in a non

desirable direction, the impact on a spacecraft's

attitude control system maybe significant. Other

issues include the effect of radiated electromagnetic emissions produced by the plasma

generated within the operating high power Hall

thruster. While this has proven to be a manageable

issue for lower power thrusters, the way in which

this scales to higher powers remains unknown.

Finally, such things as contamination of sensitive

surfaces by deposition of engine erosion products

and possible optical interference from the light

emitted by the plasma produced within the engine

must also be considered.

Discussion

In order to successfully develop the next

generation of high power Hall thrusters it is

necessary to consider each of the various topics of

consideration both separately and as they relate to

one another. With this as a basis, the individual

technological aspects of such a program can be

appropriately addressed. As previously mentioned

there is limited amount of prior data from high

power Hall thruster testing. What is noteworthy,

however, is that these data were, in general.

adequately predicted by previous data obtained

with lower power Hall thrusters. This demonstrates

the efficacy of extending the current technology in

a linear fashion which will likely provide significantly increased performance with a modest

outlay of resources. This is primarily enabled by

the very substantial body of work on Hall thrusters

at power levels of 10 kW and less.

There are physical limitations which dictate

the maximum power or power density possible with

Hall thrusters. These limitations which are not

accounted for by a scaling type of approach were

detailed previously. The most fundamental of these

physical limitations is the maximum ion current

density possible within the discharge chamber of a

NASA/TM--1999-209436 7
highpowerHall thruster.Thevalueof this limit

dependsonthepropellantandthrustertype(i.e.,

SPTorTAL),butultimatelydeterminesthesizeof

athrusterforagivenpowerlevel.Furthermore, due

totheannularnatureof Hallthrusters,asthesize

of thethrusteris increasedto permithigherpower.

theactualpowerdensityorpowerperdiameterof

thechanneldecreases. Due to this fact+as the

outerdiameterof theHallthrusteris increasedin

sizeto accommodatehigherpower,eventuallyan

ionthrusterwill theoreticallyapproachthethrustof

a Hallthrusterfora giventhrusterdiameter.It is

difficultto predictwherethesetwocurveswill

intersectbecauseof a sparcityof data,andthere

alsoremainsa significantnumberof technical

issuesto beresolvedwithregardto thedevelopment of large griddedthrusters,but for high

thrust/lowspecific impulseapplicationsit is

expectedthatthe Hall thrusterwill offerhigher

thrustdensitiesforpowerlevelswell in excessof

100kW.Ofcoursethisdoesnotconsidera multi

channel approachor possible non circular

geometrieswhichwouldsubstantiallyincreasethe

thrustdensityin comparisonto a singleclassic

annulardischargechamber.This leadsto the

conclusionthatforthrustersoptimizedto provide

highthrustwithvoltagesontheorderof 300Volts,

powersupto 100kW couldbe consideredwith

thrustersizesupto 100cmindiameter.

Therearehoweverpracticallimitationswith

regardto themaximumsizeof thrusterwhichcan

befabricated,theselimitationsarebasedonthe

availabilityof thenecessaryceramicsandrefractory metalsin thesizesneeded.Whilea detailed

assessment onthemaximumsizesofrawmaterials

were not undertaken,anecdotalexperience

indicatesthatevenat powerlevelsof 10kW one

approachesthepracticallimitsof whatis currently

available.Therearealternatestrategieswhichcan

beadopted.Usingmultiplepiecesin placeof what

hastraditionallybeena singlepieceis anobvious

approach. Theimplicationsof thiswithregardto

survivingthe mechanicalloadsassociatedwith

launchmayevenbefavorable. However,thereis a

degreeof riskassociatedwithsucha strategythat

will needtobeaddressed.

Theissueof propellantchoiceis perhapsthe

singlebiggestconsiderationwith regardto the

directionof future high power Hall thruster

development. Froma costto develop,technical

andhistoricalbasis,theexperiencecurrentlyexistingwithxenoncoupledwithfavorableperformance

makethis propellantan overwhelminglypreferablechoice.Problemswithpriceandavailability,

however,maymakethischoiceuntenablefroma

costandlogisticsstandpoint. If otherpropellants

areconsideredadditionaldevelopment is needed.

Someof this couldbe conductedwith smaller,

lowerpowerthrustersto minimizecost. The

additionalissuesthat will likely be determining

factorsforpropellantchoicesuchasenvironmental

considerationstendto be somewhatpoliticalin

natureandwill notbediscussedas partof this

paper.

Theissuesrelativeto testingwill alsohavea

verysignificantimpactonthedirectionof future

high powerHall thrusterdevelopment. This is

primarilybecausethe cost of upgradingthe

infrastructureof thevariouselectricpropulsiontest

facilitiesrequiredto testhighpowerHallthrusters

mayexceedthe costof the development effort

itself. At NASAGRC,whichhasone of the

highest pumping speed dedicated electric

propulsiontest facility, the pumpingspeedof

xenonisapproximately1 millionliterspersecond.

Fora pressureof 2x106torrthiscorrespondsto a

flowrateof 17mg/swhichat 300Voltsis onlya

5kWthruster.Inorderto conductresearchonhigh

powerHall thrustersat higherpressuresthe

implicationsof the effectof backgroundpressure

will havetobemorecompletelyunderstood.

Concluding Remarks

In summary, the implications of this study are

the following: Based on past data there are

performance benefits to be gained by increasing

thruster size. The wealth of previous experimental

data taken with thrusters of 10 kW input power and

less form an excellent basis upon which to develop

higher power Hall thrusters. There is a mass

savings in using fewer number of higher power

thrusters as compared to a larger number of smaller

thrusters to obtain a given power level. The desired

propellant is xenon, however from a cost. availability, or system perspective high power Hall

thrusters using alternate propellants may be

developed requiring a greater development effort.

There are practical limitations with regard to how

large high power Hall thrusters can be made.

Creative engineering can stretch these limits, but

ultimately development in the such things as raw

material fabrication issues may be required. The

requirements of the ground test programs

necessitated by the development of high power

hall thrusters will represent a real limitation in the

practical limit for thruster development. Finally,

all these considerations have assumed a traditional

linear or incremental development of the technology. History indicates that a technical breakthrough permitting a non linear technology jump is

possible and even likely during the development

cycle for such engines. If not, even now. Hall

thrusters with power levels up to 100 kW can be

considered which will enable missions to Mars,

NAS A/TM-- 1999-209436 8
reusablespacetugs,andotherpropulsionintensive

missionswhichhavebeenunderconsiderationfor

decades.

References

I. Oleson, S.R., Advanced Propulsion for Space

Solar Power Satellites, AIAA 99-2872

2. Gefert, L.P, Hack, K.J., Kerslake, T.W.,

Options for the Human Exploration of Mars

using Solar Electric Propulsion. proceedings of

the Space Technology and Applications

International Forum--1999.

3. Oleson, S.R., Influence of Power System

Technology on Electric Propulsion Missions,

NASA CR-195419

4. Oleson, S.R., Myers. R.M., Launch Vehicle

and Power Level Impacts of Electric GEO

Insertion, AIAA-96-2978

5. Arkhipov, B.A. et al., "Development and

Investigation of Characteristics of Increased

Power SPT Models", IEPC-93-222, Proceedings of 23-rd International Electric Propulsion

Conference, Seattle, September 1993.

8. Arkhipov, B.A. et al., "Investigation of SPT200 Operating Characteristics at Power Levels

up to 12 kW", IEPC 97-132, Proceedings of

25 'h International Electric Propulsion Conference, Cleveland, August 1997.

7. Ageev V.P., Safronov I.N., Tverdokhlebov S.O.:

Performance Characteristics and Peculiarities

of Two-Stage Accelerator with Anode Layer

Operation with Low Voltages and Magnetic

Field, Sov. Journal of Appl. Mech.&Tech.

Phys. No. 6, 1987. pp. 3-7.

8. Semenkin, A., Tverdokhlebov, S. et al., "TAL

Thruster Technology for Advanced Electric

Propulsion System". 96-a-3-26, 20 m ISTS,

May, 1996, Gifu, Japan.

9. D. Jacobson, R. Jankovsky, "Performance

Evaluation of a 50 kW Hall Thruster", ALAA

99-0457, Presented on 37a AIAA Aerospace

Sciences Meeting and Exhibit, Reno, January

1999.

10. R.S. Jankovsky, C. McLean, J. McVey,

"Preliminary Evaluation of a 10kW Hall

Thruster", AIAA-99-0456, Presented on 37t"

AIAA Aerospace Sciences Meeting and

Exhibit, Reno, January 1999

11. Melikov, I. et al., Soviet Journal of Technical

Physics, 1974, Vol. 44, No3.

12. Zhurin, V. et al., "Physics of Closed Drift

Thruster", IEPC-97-191. 25'h IEPC, August,

1997, Cleveland.

13. Arkhipov, B. et al. "Extending the Range of

SPT Operation: Development Status of 300

and 4500 W Thruster". AIAA-Paper.

Presented on 32 ndJoint Propulsion Conference,

Orlando, 1996

14. Gavryshin, V.M., Kim, V. et al., "Physical and

Technical Bases of the Modern SPT

Development", IEPC 95-38, Proceedings of

24 _h International Electric Propulsion

Conference, Moscow, 1995.

15. Sankovic. J., Hamley, J., and Haag, T.,

"Performance evaluation of the Russian SPT100 Thruster at NASA LeRC," AIAA-93-094,

September. 1993.

16. Bugrova, A. et al., "Scaling Laws for Integral

Petformances of SPT", Soviet Journal of

Technical Physics. 1991. Vol. 61. No. 6,

pp. 35-51.

17. Bugrova, A., Kim, V.. "Modern State of

Physical Investigation in the Accelerators with

Closed Electron Drift and Extended

Acceleration Zone "--Paper in the book:

Plasma.

18. Marrese, C. et al., "D-IO0 Pelformance and

Plume Characterization on Krypton ",

AIAA-96-2969, 32 _J AIAA JPC, July, 1996,

Lake Buena Vista, FL.

19. Manzella, D. and Sankovic, J "Hall Thruster

ion beam Characterization," AIAA-95-2927.

July, 1995.

NASA/TM--1999-209436 9
Table1:Performance ofpreviouslydevelopedlaboratorymodelhighpowerHallthrusters

Thrustertype

Averagediameter,

mm

Isprange,s

Inputpower,kW

Thrust.mN

SPT-290

Ref.5

250

1500-3000

12-30

T-220

Ref.10

188

1500-2400

5-11

TALTM-50

Ref.9

200

1500-3300

10-25

TAL-200[Bi]

Ref.7

200

2000-5200

10-34

SPT-200

Ref.6

175

1500-3000

6-11

1500 524 1114 1130 498

Efficiency 0.7 0.62 0.66 0.67 0.63

Table2:Comparisonofalternateelemental propellants

Element

Radon

Molecular

Weight

FirstIonization

Potential, eV

%Thrust

ofXenon

222 10.7 130

% Isp

of Xenon

77

Bismuth 208.98 7.3 126 79

Lead 207.19 7.4 126 80

Mercury 200.59 10.4 124 81

Cesium 132.905 3.9 101 99

Xenon 131.30 12.1 100 100

Krypton 83.80 14.0 80 125

Argon 39.948 15.8 55 181

t500

Z

E

.,d'

[

[,.

1000

50O

tl,

n

* i

AQ

[]

[]

• U •

m

0

0

[]

m _ •

• SPT-200

•'T-220

• SPT-290

i I_TM-50

I

• TAL-200

10O00 20000 30000 40000

Power, Watts

Figure 1: Thrust versus power of previously tested high power Hall thusters

5O00O

NASA/TM-- 1999-209436 10
6O

5O

40

030

a.

t"

2O

10

0

o SPT, Isp=2500 s

" TAL, I,sp=2500 s

[] SPT, ,l_p=2000 s

1

• TAL, Isp=2000 s

+ TAL, Isp=3000 s

• TAL, Isp=1700 s

+

100 200 300 400

Average Diameter, mm

Figure 2: Projected Effect of Input Power and Specific Impulse on Thruster Size

5O0

NASA/TM-- 1999-209436 11
m

e,,

_E

.o

4500

4000

3500 -

3000 "

2500 -

2000

1500 -

1000 -

500 -

o o

o J-_ 1.1-11_

0 i_ ..f._ _ .lllfJ

f/1-

jt_

ft I Io SPT

• TAL

0 .............. T" _ - "

0 0.05 0.1 0.15 0.2 0.25

m/D, mg/s mm

Figure 3: The ratio of mass flow rate divided by average thruster diameter versus specific impulse

-i

0.3

5O

,,r 30

E 20

4O

10

, J

0 5 10

J I

J

J

J

i i

15 20 25 30 35 40 45 50

Thruster Power, kW (|s[_/,5_ s)

Figure 4: Projected thruster mass versus power for flight and engineering model thrusters

NASA/TM--1999-209436 12

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1. AGENCY USE ONLY (Leave blank) 2. REPORT DATE 3. REPORT TYPE AND DATES COVERED

December 1999 Technical Memorandum

4. TITLE AND SUBTITLE 5. FUNDING NUMBERS

High Power Hall Thrusters

6. AUTHOR(S)

Robert Jankovsky, Sergey Tverdokhlebov, and David Manzella

7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration

John H. Glenn Research Center at Lewis Field

Cleveland, Ohio 44135-3191

9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration

Washington. DC 20546-0001

WU-632- IB- 1B-00

8. PERFORMING ORGANIZATION

REPORT NUMBER

E-11912

10. SPONSORING/MONITORING

AGENCY REPORT NUMBER

NASA TM--1999-209436

AIAA 99-2949

11. SUPPLEMENTARY NOTES

Prepared for the 35th Joint Propulsion Conference and Exhibit cosponsored by AIAA, ASME, SAE. and ASEE,

Los Angeles, California, June 20-24, 1999. Robert Jankovsky. NASA Glenn Research Center, Cleveland, Ohio, Sergey

Tverdokhlebov. TsNIIMASH Export. Korolev. Moscow. Russia and David Manzella, Dynacs Engineering Company,

Inc., Brook Park. Ohio. Responsible person. Robert S. Jankovsky, organization code 5430, (216) 977-7515.

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12b. DISTRIBUTION CODE

13. ABSTRACT (Maximum 200 worde)

The development of Hall thrusters with powers ranging from tens of kilowatts to in excess of one hundred kilowatts is

considered based on renewed interest in high power, high thrust electric propulsion applications. An approach to develop

such thrusters based on previous experience is discussed. It is shown that the previous experimental data taken with

thrusters of 10 kW input power and less can be used. Potential mass savings due to the design of high power Hall thrusters

are discussed. Both xenon and alternate thruster propellant are considered, as are technological issues that will challenge

the design of high power Hall thrusters. Finally, the implications of such a development effort with regard to ground

testing and spacecraft integration issues are discussed.

14. SUBJECT TERMS

Hall thruster

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