John E. Foster, Aimee Hubble, and Sarah Nowak-Gucker

The University of Michigan, Ann Arbor, Michigan 

Chris Davis, Peter Peterson, Eric Viges, and Dave Chen

ElectroDynamic Applications, Inc., Ann Arbor, Michigan

Simulated Beam Extraction Performance

Characterization of a 50-cm Ion

Thruster Discharge

NASA/CR—2013-217828

June 2013

AIAA–2012–3795
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John E. Foster, Aimee Hubble, and Sarah Nowak-Gucker

The University of Michigan, Ann Arbor, Michigan 

Chris Davis, Peter Peterson, Eric Viges, and Dave Chen

ElectroDynamic Applications, Inc., Ann Arbor, Michigan

Simulated Beam Extraction Performance

Characterization of a 50-cm Ion

Thruster Discharge

NASA/CR—2013-217828

June 2013

AIAA–2012–3795

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Prepared under Contracts NNX10CF59P and NNX11CC56C

Prepared for the

48th Joint Propulsion Conference and Exhibit

cosponsored by the AIAA, ASME, SAE, and ASEE

Atlanta, Georgia, July 30–August 1, 2012
Acknowledgments

This work is supported by NASA under the STTR program, contract numbers: NNX10CF59P and NNX11CC56C.

Available from

NASA Center for Aerospace Information

7115 Standard Drive

Hanover, MD 21076–1320

National Technical Information Service

5301 Shawnee Road

Alexandria, VA 22312

Available electronically at http://www.sti.nasa.gov

Level of Review: This material has been technically reviewed by expert reviewer(s).

This report contains preliminary fi ndings, 

subject to revision as analysis proceeds.

I. Introduction

DVANCES in high power, photovoltaic 

technology has enabled the possibility 

for reasonably sized, high specific power, 

high power, solar arrays. Indeed, specific 

power estimates for thin film solar arrays 

is an order of magnitude higher than that 

of the present state of the art, which 

currently stands at around a few hundred 

Watts per kg.1 Thin film solar arrays have 

projected specific powers of over 4000 

W/kg.2 These values even exceed those of 

combined solar concentrator and solar 

array systems currently under 

development (~130 W/kg) such the Fast 

Access to Space (FAST) system.3 Even 

so, array technologies such as the FAST 

system are mission enabling for both near 

and outer planet missions (see Figure 1). 

At such high specific powers, power 

levels ranging from 50 to several hundred 

kW are feasible for communication 

satellites in low volume, low mass packages, thereby greatly increasing their capability and overall lifetime. 

Coupled with electric propulsion, gridded ion thruster technology in particular, this power technology can be 

mission enabling for a wide range of missions ranging from ambitious near Earth NASA missions to those missions 

involving other customers such as DOD and commercial satellite interests. Station keeping for space solar power 

stations is also another mission that could be enabled by the marriage of high power, lightweight solar arrays and 

electric propulsion.

The appeal of the gridded ion thruster stems from the engine’s overall high efficiency, over 70% for high 

specific impulse applications, and long operational life.4,5 At present, the most advanced and mature gridded ion 

thruster technology is that embodied in the 7-kW NEXT ion thruster. But currently there is no single, high TRL 

level (>4) gridded ion thruster capable of processing power beyond 7-kW. It is envisioned that in the near-term, 

there will be a need for single thrusters with the capacity to process greater than 10 kW.6 This need is commensurate 

with power expected to be available and the associated missions that this combination enables. In general, single 

thruster systems capable of high-power operation are attractive in that they reduce system complexity as well as 

system mass. The aim of the effort summarized herein seeks to design and fabricate a partial conic geometry ion 

thruster discharge chamber with an equivalent beam area of a 50-cm-diameter cylindrical ion thruster with the 

capacity to fill the 7 to 25-kW void that currently exists for gridded ion thrusters. Again, such a thruster fulfills a 

need in that it fills the power gap void, allowing for the development of thruster technologies that enable NASA to 

respond rapidly to those new and emerging missions that may require high power as enabled by recent advances in 

photovoltaic arrays as well as be responsive to non-NASA customers such as DOD and commercial interests for 

Near-Earth station keeping and orbital transfer applications and beamed solar power. The overall effort will advance 

the TRL level of the 50-cm thruster by understanding and optimizing the performance of the discharge chamber. By 

boosting the TRL level on the discharge chamber, realization of a 7-25 kW gridded ion thruster in the mid-term 

becomes straightforward with subsequent development of the ion optics and the power-processing console to follow. 

The discharge chamber operating characteristics in part determine the requirements for the power processing unit as 

well as the ion optics design so its advancement is critical for the realization of a high-power thruster in the midterm.

A. Historical Perspective

During the 1980s, NASA investigated the performance of a 50-cm ion thruster featuring a conventional, “coffee 

can” geometry. The 50-cm test article was meant to be a test bed to the development of high power ion thrusters. At 

A 

Figure 1. FAST solar array enables both near Earth and 

outer planet missions (graphic courtesy of DARPA). 

NASA/CR—2013-217828 2
the time the projected power range for the single 50-cm thruster was up to 20 kW.7 In those tests, while a number of 

magnetic circuit configurations were investigated, the magnetic circuit was not optimized. Despite this, the large 

thruster operated at discharge losses typically below 150 W/A. Discharge stability, however, was found to be quite 

poor. Discharge stability is determined in large part by the geometry of the magnetic circuit.8 The 50-cm thruster 

with 50-cm optics operated with a nominal discharge voltage of 30 V, which is above the sputter threshold of grid 

construction materials such as molybdenum. Additionally, thermal loading on thruster magnets could not be 

adequately addressed in this effort for two reasons: 1) magnets were uncovered and as a consequence some of which

suffered irreversible field loss due to heat loads which in turn had a deleterious effect on performance; and 2) 

thermocouples were not used to assess magnet temperature.

Between the late 1980s and the year 2000, there was essentially a hiatus in the development of larger thrusters. 

The 50-cm thruster did not advance beyond the aforementioned demonstration effort but the effort did result in a set 

of 50- cm ion optics being fabricated. Patterson in 2000 investigated the use of three individual sets of ion optics 

mated to a single large discharge chamber as a means to achieving higher power.9 This approach exploited the 

heritage and high perveance capacity of 30-cm ion optics. In this case a 76-cm-diameter discharge chamber was 

used. One issue with this approach was non-uniformity of plasma across the individual ion optics sets. Presumably, 

this issue could be minimized by further optimization of the magnetic circuit but in its present configuration, the 

plasma non-uniformity cannot completely be eliminated because of the inherent asymmetry (ion optics sets were off 

axis). Later, 50-cm optics (derived from the 50-cm thruster effort) was also investigated for high voltage beam 

extraction studies using the 76-cm discharge chamber to supply beam ions.7 High-power (i.e., 20-kW-class) ion 

thrusters were also investigated during the JIMO program. Parallel efforts at NASA Glenn Research Center (GRC) 

and the Jet Propulsion Laboratory (JPL) developed large-area thrusters essentially to TRL 4. These very high 

specific impulse thrusters (>5000 s) are not suitable for near Earth applications owing to the very low thrust-topower ratios. Rather, they are more suitable for outer planet missions. Development of these thrusters to flight-like 

status is expected to require many years of additional component level testing to increase the TRL level of the 

discharge cathode, the ion optics and the high-voltage isolation.

For mid-term applications, particularly those involving near-Earth missions, it has been suggested that 50-cm 

variants be investigated to provide NASA with a more powerful thruster option beyond NEXT.6 Such a thruster 

would operate at considerably higher beam currents and thrust than the NEXT thruster (~ 4 x beam current, ~3 x 

thrust).6 As identified in the Patterson report, a 50-cm device could be advanced to a TRL 6 in just six years 

provided key technological issues are thoughtfully addressed first, one of the goals of this effort.

Three technology areas must be addressed to advance the state of the art for the 50-cm thruster. These include: 

1) the discharge chamber magnetic circuit; 2) the cathode and neutralizer; 3) and high-perveance 50-cm ion optics. 

We report herein the status of the development of this engine. In this effort a new discharge chamber for a 50-

cm ion thruster was designed, fabricated and tested under simulated beam extraction conditions. The thruster 

will be capable of generating an ion beam of 14 A at a 25 kW. This power range addresses the single–thruster power 

gap that currently exists.8 This effort starts with simulated beam extraction to characterize and optimize the 

discharge chamber, but will culminate in a beam extraction test for a “full-up” performance test to be compared with 

the simulated beam extraction tests for final validation. 

B. Unique Contributions of Task

The discharge chamber design developed in this effort differs from the explorative work done in the late 1980s 

in two key aspects. First, a numerical magnetic solver was used to design the initial magnetic circuit. In this current 

investigation, extensive internal plasma diagnostics are used along with the numerical solver to optimize the 

magnetic circuit design. Design and optimization of the source in this self-consistent manner allows us to address 

plasma uniformity at the exit plane, which determines the ion optics design and discharge efficiency, which in turn 

determines power requirements and ultimately contributes to the overall efficiency calculation. Plasma collection at 

the magnetic rings is measured to assess thermal loading of at the magnets. Using plasma diagnostics and 

thermocouples will allow for an a priori mapping of the operating envelope.

The second difference is that this effort addresses mass reductions up front. A partial conic design reduces 

mass in contrast to the conventional cylindrical geometries investigated in the 80s. The partial conic geometry not 

only has flight heritage (NSTAR to the Dawn missions) but it is also the embodiment of the state of the art as 

manifested in the NEXT thruster. Additionally, the use of smaller, higher-temperature, higher-energy product 

magnets is expected to further reduce thruster mass. Altogether, this effort focused on the design and fabrication of 

a new type of 50-cm ion thruster discharge chamber – one designed to be lighter than previous devices, more stable 

over a broad operating range, and efficient.

NASA/CR—2013-217828 3
II. Technical Approach and Design Considerations

 At the heart of the discharge chamber design 

process is the magnetic circuit. The magnetic 

circuit was designed using the commercial 

magnetostatic solver Maxwell 3D. A 45-degree 

slice of the discharge chamber was modeled to 

reduce simulation time. Several rules of thumb 

were used in the design of the magnetic circuit. 

Following Wirtz et al, an even number of magnet 

rings was chosen, as it assures a magnetic 

minimum on axis and thus avoids the generation 

of peaked discharge profile derived from over 

confinement of primary electrons.10

The initial discharge chamber geometry 

dictated the position and number of magnet rings 

necessary to satisfy the aforementioned criteria 

governing the magnetic circuit. The rare earth 

magnets which make up the magnetic circuit 

were high temperature, rare earth magnets.

The discharge chamber consists of two 

sections: a conical section and a cylindrical 

section. The overall length of the discharge 

chamber could be extended by adding cylindrical 

section inserts, which were each approximately 

20 percent the baseline thruster length. Two such 

inserts were fabricated. The inserts were used to 

understand the sensitivity of propellant utilization 

to chamber length. The two inserts run together 

required the addition of another magnet ring to 

prevent the formation of lossy virtual cusps at the 

wall. The magnetic circuit for each discharge 

chamber configuration satisfied the circuit rules 

outlined earlier. Figure 2 depicts the calculated 

baseline discharge chamber configuration with 6 magnet rings. The fidelity of the calculated field was validated 

with internal point by point Gauss meter measurements.

The magnet rings were covered using electrically isolated, stainless steel covers. Each magnetic ring 

actually consisted of ring segments—six per magnet ring. The electrodes were electrically isolated from anode 

potential using a combination of ceramic bushings and thin mica washers. The purpose of the ring electrodes was to 

measure current collected at each magnet ring. Past studies have indicated a strong correlation between relative 

magnitude of collected current at the magnet rings and available ion current in the adjacent volume.11 Knowing the 

spatial distribution of the collected current therefore lends a great deal of insight into the location of the centroid of 

the bulk plasma. Such insight is used as a guiding tool during the optimization studies. The ring electrodes also 

serve as a physical shield or cover, protecting the magnet rings from excessive heat flux associated with the very 

localized electron collection. The engine itself is depicted in Figure 3

The performance testing with simulated beam extraction allows for optimizing engine performance at the 

desired operating conditions. Targeted operating conditions are depicted in Table 1. Table 2 represents the idealized 

upper limit for performance. Here thrust losses such as plume divergence, double to single ratio, and neutralizer 

power is not incorporated in the calculations shown in the Table. These conditions represent engine throttling at 

constant beam current. A chief goal of this effort was to maximize thrust to power. This can be achieved in part by 

reducing the discharge losses. In this engine, optimization efforts target discharge losses less than 120 W/A. The 

highlighted low beam voltage point represents a very high thrust to power operating condition. Demonstration of 

operation at such conditions even with limited life is of interest for quicker orbit transfer applications. The 1800 V 

operating conditions refers to a high-powered, high specific impulse operating condition, compatible with tug-like 

operations for Mars for example. 

Figure 2. Magnetic circuit layout of baseline engine 

configuration. Aspect ratio not to scale.

Figure 3. Assembled 50 cm hardware minus grids.

NASA/CR—2013-217828 4
Vbeam

[V]

Ibeam

[A]

T

[ mN]

Isp

[sec]

Pin

[W]

T/P 

[mN/kW]

Total 

Efficiency

1800 14 950 4700 27000 35 0.81

1550 14 900 4400 23500 38 0.81

1400 14 830 4100 21000 40 0.80

1200 14 770 3800 18000 43 0.79

1020 14 710 3500 15800 45 0.78

950 14 680 3400 14600 47 0.77

850 14 650 3200 13400 48 0.76

700 14 600 2900 11300 52 0.75

Table 1. Target operation parameters of the HPHTion

III.Test set up

The discharge chamber was tested over a wide range of conditions. 

To assess the performance of the engine and quantify changes to 

plasma conditions, a number of diagnostics beyond the general power 

console response were implemented. In this section, the general 

description of the test facility, power console and diagnostics is given. 

A. Vacuum facility and feed system

Simulated beam extraction testing took place at the University of 

Michigan PEPL Large Vacuum Test Facility (LVTF). The LVTF 

(Figure 4) is 9 m long and 6 m in diameter. This facility is evacuated 

by seven LN2-cooled CVI TM1200 reentrant (nude) cryopumps that 

give the facility an overall pumping speed of 245,000 l/s on xenon and 

a base pressure of <2x10-7 Torr and the LVTF can maintain a pressure 

in the 10-6 Torr range (xenon pressure) during 

the operation of a 20 kW ion thruster with a 

beam current of 18 A. A louvered 1.8 m by 1.8 

m graphite beam dump is located on the center 

of the endcap downstream of the thruster to 

minimize the deposition of back-sputtered 

material from the bare tank wall. 

 In addition, the LVTF is equipped with a 

high capacity turbo pump that allows 

regeneration to occur at background facility 

pressure of approximately 10-4 Torr. The PEPL 

facility is also equipped to operate various inspace propulsion systems with dedicated power 

supply racks and mass flow systems for ion 

thrusters up 8 kW and Hall thrusters up to 100 

kW. The existing high purity xenon flow 

system was utilized for these tests. Xenon of 

purity 99.999 was used for these tests. For all 

tests, the background pressure was less than 1.5 

x 10-5 Torr

B. Power console and diagnostics

Simulated beam extraction testing setup follows from that proposed by Brophy.12 Laboratory power supplies 

were used to operate engine and power the diagnostics. The main discharge supply was a 60 A, 40 V power supply. 

Uncertainty in discharge voltage was approximately +/- 1V. Discharge voltage oscillations were monitored using an 

oscilloscope. The anode return current was comprised of seven components as illustrated in Figure 5. These 

included the currents collected at the 6 magnet rings and the current collected between the magnet rings—the anode 

shell current. The collected current flowed through magneto resistive current channels—each of which were 

monitored using a conventional A/D convertor and computer which allowed for visual strip chart monitoring of the 

Figure 4. PEPL LVTF facility.

Figure 5. Anode current electrode and measurement 

scheme

Figure 6. Ion collection grid with array of fixed plasma probes 

NASA/CR—2013-217828 5
current distribution to the anode. In this investigation, 

the discharge current did not exceed 46 A—a selfimposed limit placed on the 50 a cathode. Resistance 

between ring electrodes and the anode shell was high 

typically ~Giga-Ohms as indicated by “hi-potting” at 500 

V- 1000 V, indicating very good isolation. 

The main discharge chamber was terminated with a 

stainless steel grid with an effective open area fraction of 

32%. This grid was used to estimate the extractable ion 

current at a given operation condition. The grid to anode 

gap was approximately 2 mm. The grid was biased 

relative to the cathode at a approximately -20 V, which 

was sufficient for ion saturation. In order to assess 

discharge uniformity, the grid was fitted across its 

diameter with twenty-five 6 mm diameter molybdenum 

button probes. This grid is depicted in Figure 6. These 

electrically isolated probes, spaced 20 mm apart, were 

each biased -20 below cathode potential. The current 

signal from each probe was measured from precision 

100-Ohm resistor and strip 

chart monitored using an 

analog to digital converter. 

Acquisition rates were 

typically ~ 1-2 Hz. In addition 

to the grid probes, a single 

Langmuir probe was used to 

measure plasma properties in 

the bulk of the discharge. The 

probe, located in the middle of 

the cylindrical section, 

extended radially 16.5 cm into 

the discharge chamber. The 

magnetic field at this point was 

approximately 1 G, indicating 

an essentially field free bulk 

region. 

Select magnet rings were 

monitored using type K 

thermocouples. In this work, 

the four most downstream rings were monitored as these rings collected the majority of the discharge current. 

Thermocouples were in physical contact with magnets that comprised the various magnet rings for an accurate 

assessment of heat loading. Temperature measurements were acquired from the thermocouples using a hand-held 

electrically isolated readout. 

IV.General Results

In order to understand the general operating behavior of the discharge chamber was operated in simulated beam 

extraction mode. Simulated beam extraction represents a cost effective method for understanding discharge 

operation. Low voltage operation associated with simulated beam extraction, makes the discharge amenable to the 

integration of plasma diagnostics as well such as the anode ring collectors and grid ion probes. Simulated beam 

extraction tests carried out here follows the work of Brophy. In this specific case, a single grid terminates the 

discharge chamber exit plane. Under simulated beam extraction conditions, those ions that do not escape through 

the apertures are neutralized on the upstream surface of the single grid. These neutralized ions will reenter the 

discharge chamber, giving rise to effectively a larger mass flow rate. To simulate neutral density conditions similar 

to the case with actual beam extraction, the flow must be throttled back to account for this additional flow source. 

Figure 6. Ion collection grid with array of fixed 

plasma probes to assess discharge uniformity.

Figure 7. Discharge efficiency curves at fixed discharge current.

NASA/CR—2013-217828 6
The relations for the simulated current, the ion transparency to the single grid and the simulated propellant 

efficiency are therefore:

Jbeam−sim = Jgrid ⋅φi

where is the grid transparency and is the total ion flux incident on the grid. Grid transparency may be also 

be determined by: 

φi = Jgrid

Jgrid + Jbeam

The simulated utilization can therefore be calculated: 

ηm−sim = φi ⋅ Jgrid

m + Jgrid

1−φi

φi

#

$

% &

'

(

along with discharge losses: 

η D = JDVD

Jbeam−sim

In this manner, thruster operation will be simulated. Studies have shown that this approach replicates actual thruster 

operation very well.13

A. Discharge performance

In order to assess the discharge efficiency, the discharge chamber’s simulated discharge losses and propellant 

utilization were determined from grid current measurements, discharge input power, and mass flow rate. The 

performance curves were acquired using two methods: 1) vary the flow at fixed discharge current 2) vary discharge 

parameters such as main and cathode flow rates as well as the discharge current at fixed simulated beam current. 

In these investigations, discharge voltage varied between 24 and 29 volts nominally. Some operating points at 

discharge voltages as high as 30 V were also acquired to map out the region beyond the “knee.” Figure 7 is the

performance plot obtained by fixing the discharge current and varying the flow. When this is done, the simulated 

beam current increases as does the utilization. Such plots indicate the simulated beam current range for a given 

input discharge power. The 37 A and 42.2 A cases follow expected performance trends with knees at just over 0.8 

and 0.87 respectively. The 32 A case is somewhat puzzling in that the discharge losses appear to be invariant to 

changes in flow over a broad range. Here, with changes in flow rate, the discharge power and grid current increase 

at the same rate. In all cases, these data sets suggest low discharge losses ranging from ~125 W/A to as low as 115 

W/A for the baseline configuration. The discharge was also operated at 44 A, which resulted in even lower 

discharge losses—close to 110 W/A at 0.80 utilization. For comparison, the 37.1 A operating condition case 

approached this discharge efficiency at the relatively low utilization of 0.67. Since during normal thruster operation, 

the beam current is typically fixed, it is of interest to plot performance at fixed extractable grid current. Figure 8 

illustrates the behavior of the discharge efficiency at fixed simulated beam current. In all cases, the cathode 

operated in essentially a quiet mode as inferred from peak to peak voltage oscillations on the discharge voltage 

(oscillations ~5 V average), oscillations in probe current at the grid, and visual observation of the discharge. 

Simulated beam currents up to nearly 12 A were achievable with the existing configuration. Higher beam currents 

require operation at higher discharge currents. The existing discharge cathode is limited to around 50 A of emission 

current. A somewhat larger cathode emission current will be required to achieve the 14 A beam case (~55-60 A). 

These plots also indicate the relatively low discharge losses—in this case occurring less than 120 W/A. At the 11.9 

A condition, losses were around 110 W/A at a utilization of nearly 0.9. These discharge losses compare with 

optimized engines such as NEXT , which operates at around 125 W/A and 0.89 utilization at full power.14 In any 

event, the performance values inferred from simulated beam extraction at the lower beam currents compare well 

with the design target of ~110 W/A at utilizations of around 0.92 for the 14 A beam case. The data suggests that the 

14 A performance target is achievable based as the performance curve knee tends to shift with increasing discharge 

current towards higher utilization. The low discharge losses overall appear to be attributed to three factors: 1) Large

discharge chamber translates into longer characteristic diffusion times. 2) Optimal magnetic circuit featuring 6 rings 

minimizes losses 3) Discharge chamber is masked down by 15 percent.

φi Jgrid

NASA/CR—2013-217828 7
A. Discharge Length Sensitivity 

One parameter that controls the 

ionization efficiency of the discharge 

chamber is the physical length to 

diameter ratio of the discharge 

chamber. This parameter in 

conjunction with the grids 

determines the neutral residence time 

and thus the neutral containment. In 

this work, this parameter was varied 

by adding a cylindrical insert, which 

was approximately 20 percent of the 

total baseline length of the discharge 

chamber. Increases in chamber 

length appreciably longer than 20 

percent would require additional 

modification to the magnetic circuit 

to prevent the formation of lossy 

virtual cusps. Fig ure 7 and 8 show 

variations in discharge performance 

with the insert. On one hand, the insert does appear to improve discharge losses at the low utilization points as 

indicated in Figure 7, but overall as indicated in Figure 8, the addition of the insert apparently degraded performance 

somewhat. This is attributed in part to the combination of increased surface area and reduced magnetic field in the 

region of the insert. With the exception of low utilization operating conditions, as utilization gets higher the reduced 

magnetic field and additional surface area presumably offsets any performance gains associated with improved 

neutral containment with the longer chamber. 

B. Plasma Uniformity

The plasma density profile across the ion optics directly impacts overall engine performance. The curvature of 

the plasma meniscus, whose shape is a function of the plasma density and the beam voltage, in part determines the 

divergence angle of each beamlet thus impacting thrust efficiency. Typically at large off axis radii, reduced plasma 

density results in poorly focused beamlets. While it is possible to “engineer” this problem away by varying aperture 

hole size and perhaps chaffering the apertures as a function of radius, the usual situation—which is also the most 

cost effective is to implement grids with fixed hole sizes and spacing that don't vary with radius. Such optics in 

principle, unlike an engineered aperture profile, can run over a variety of operating conditions. One simply maps 

out the envelop of operation; that is, one determines the regimes of operation at both high and low power levels 

where impingement and divergence are at acceptable levels, which are usually dictated by lifetime requirements—

with margin. In this work, the ion density profile in the plane of the optics was assessed using 25 fixed button probes 

separated 2 cm apart mounted along a diameter, with active area facing the cathode. These probes were biased in 

the ion saturation regime at 20 Volts below cathode potential. If one assumes that the plasma electron temperature 

is approximately constant at least over the diameter of the unmasked region of the discharge chamber, then the 

current density profile also represents the plasma density profile. This assumption is justified in that interrogated 

area is just downstream of the field free region of the bulk discharge which is typical free of substantial gradients. 

This is the usual assumption also made with Faraday probe sweeps in the plume as well. The assumption of a flat 

electron temperature profile in the bulk region within the discharge chamber is consistent with actual spatial 

measurements of plasma properties made in the bulk in similar discharge studies.13 Figure 9 depicts discrete current 

profile data acquired using the button probes at a representative operating point. From data such discrete data, an 

estimate of the expected in beam flatness can be made. As the probes represent essentially 25 equally spaced 

measurements across the plume, the average over these 25 current “bins” will yield an average beam current density. 

The flatness parameter, which is defined as the average current density to the peak current density, can therefore be 

calculated. The peak current density is simply the maximum value over the 25 current “bins.” The profile in Figure 

9 represents a flatness of 0.80. In general, over all operating conditions investigated, the flatness parameter ranged 

from 0.7 to 0.8. The lower flatness was observed at those conditions of reduced flow rate. Table 2 yields simulated 

beam extraction conditions and observed flatness parameters for representative operating conditions. Simulated 

propellant utilization at both operating conditions was approximately 0.9. Both conditions demonstrate reasonable 

Figure 8. Discharge efficiency curves at fixed beam current.

NASA/CR—2013-217828 8
flatness even at such high utilizations. The 

flatness parameter numbers reported here are 

conservative owing to the discrete nature of 

the data, which for a finite set of probes 

would tend to overly represent peaked 

features.

Discharge Voltage, V Discharge Current, A Sim. Beam Current, A Flatness

27.2 37.1 8.7 .7

28.0 42.2 10.4 .7

Table 2: Simulated beam extraction conditions and observed flatness parameters.

C. Current distribution at magnet rings

In additional to mapping general performance of the discharge chamber and plasma uniformity, overall engine 

operating characteristics and behavior is intimately tied to cathode operating mode and the magnetic circuit. The 

majority of the discharge current is collected at the various magnet rings. The current distribution in turn determines 

the spatial distribution of the plasma in the discharge chamber. This follows from the fact that in ring cusp engines, 

peripheral ionization takes place at the magnet rings. Collection of plasma current at and subsequent diffusion of 

plasma from these localized plasma sources determine everything from the bulk plasma spatial distribution to 

discharge stability to the thermal limitations of the device. In this work, ring stainless steel electrodes covered each 

magnet ring. These electrodes allowed for monitoring of the current distribution as a function of operating condition. 

Current to anode surfaces between magnet rings—denoted the “shell”—was also monitored in this work. Figure 10a

illustrates the behavior of the currents at the various rings as at three different propellant utilizations but at a fixed 

simulated beam current of 11.9 A. The currents are normalized to the total discharge current. As can be seen here, 

the majority of the current collection occurs at the four most downstream magnet rings, giving the current collection 

profile a parabolic shape. The shell collects 13-15 percent of the discharge current and constitutes a higher 

collection fraction at the highest utilization. 

Owing to the differences in collection area, the 

current density profile differs somewhat from 

the raw current profile. Here, the behavior of 

the current density is estimated by dividing the 

collected current at a given ring by the 

circumference of that particular magnet ring 

electrode. A plot of this parameter is shown in 

Figure 10 b. It should be pointed out that this

parameter is a rough estimate as the actual 

collection area depends on the spatial variations 

in the magnetic field at the cusp. This current 

density parameter plot suggests current density 

is higher in the conical section and drops off 

towards the cylindrical section. As can be seen 

in Figure 10, the current collection profile does 

not change appreciably with increasing 

utilization. In all cases, ring 3 appears to be the 

dominant ring from both a raw collection and 

current density standpoint. The profiles 

presented in Figure 10 were prevalent over 

Figure 9. Ion current profile across simulated optics. 

Flatness =0.8. 

Figure 10a. Variation in anode current distribution with 

utilization.

NASA/CR—2013-217828 9
most of the operating conditions investigated. 

In many respects, this feature is desirable from 

a stability standpoint. If the current collection 

profile is relatively constant over a broad 

range of operating conditions, then the 

engine’s operating attributes such as beam 

flatness and thermal loading should be 

predictable to first order. Wildly varying 

current distributions would otherwise lay 

constraints on the operating envelop as well as 

require detailed testing to understand impact at 

a range of different operating points. 

It is interesting to note that as the engine 

flow was leaned out substantially; definite 

structure in the plasma profile across the exit 

plane was observable. This structure appears 

to coincide with the magnet ring locations 

particularly in the conical section. As the 

rings are a source of plasma production, it is 

not surprising at the lower flow rates that this 

is apparently observable. The data does appear to suggest that the ultimate beam profile is due to the supposition to 

the individual plasma sources at the various rings and the cathode plume. Apparently under certain conditions, this 

discrete structure persists all the way up to the exit plane of the engine. For large engines, lean flow operation may 

result in some additional structure in the ultimate exit plane profile.

D. Plasma data

A single Langmuir probe was used to assess bulk plasma properties in the cylindrical section of the discharge 

chamber. In general, the plasma in this section is determined in part by rings 5 and 6. This region was chosen 

because it is representative of plasma 

conditions just upstream of the ion 

optics. Thin sheath analysis was used 

to extract plasma density and electron 

temperature from the IV characteristics. 

The electron temperature was 

determined from the electron-retarding 

region after total probe current was 

corrected by subtracting the ion current. 

To avoid perturbing the plasma, the 

probe was not driven to the plasma 

potential to evaluate the density. 

Rather, only the electron retarding 

region and the ion saturation region was 

captured over a sweep. Ion saturation 

current is related to plasma density 

through the following thin sheath 

relation:

€

Isi = 0.61⋅ e ⋅ no ⋅ kTe

mi

⋅ Ap

A typical IV characteristic is shown in Figure 11. Table 3 lists the plasma density and electron temperature 

variations in the cylindrical section at 11.9 A of simulated beam current at different propellant utilizations. The 

electron temperature tended to increase with increasing utilization. This follows as discharge voltage increases with 

utilization, which in turn increases primary electron energy. Additionally, reduction in main flow at approximately 

fixed power results in a reduction of neutral density. In electropositive discharges such as this, electron temperature 

increases with decreasing neutral density. These trends are therefore consistent with associated variation in 

Figure 11. Typical IV characteristic of cylindrical section plasma

Figure 10b. Variations in current density parameter with 

utilization. 

NASA/CR—2013-217828 10
discharge input parameters. The plasma density on the other hand remains relatively constant over the variation in 

utilization at least within experimental uncertainty in this measurement. As shown here, the electron density is fairly 

high, nearly 1012/cc. The relative constancy of the plasma density with increasing utilization is consistent with the 

relative invariance in current collected at the cylindrical section magnet rings. Presumably, increases in electron 

temperature are not sufficient to give rise to appreciable rises in the electron density. Likely this is due to reduced 

main flow at the higher utilizations. Even though the plasma electron temperature increases, the ionization rate is 

directly proportional to the neutral pressure, which is decreased in order to achieve higher utilization. These two 

effects apparently offset each other. 

Utilization Ne, /cc Te, eV

0.75 6.35 x 1011 4.19

0.83 6.15 x 1011 4.94

0.93 6.36 x 1011 5.84

Table 3: Langmuir Probe Data—11. 9 A Simulated Beam Current.

E. Thermal Data

The temperature of the four most 

downstream magnet rings was monitored as 

a function of operating condition using type 

K thermocouples. These rings were of 

interest primarily because the majority of 

the plasma current was collected there and 

therefore these rings were expected to 

experience the greatest thermal loads. The 

magnets utilized in this study had an upper 

temperature operating limit of 600 C. In 

this work, the magnet rings monitored 

never exceeded ~350 C, indicating at least 

over the regime characterized and reported 

herein, the magnets were always well below 

magnet thermal limits. Figure 12 

illustrates typical thermal behavior; in this 

case at the highest discharge current 

condition investigated corresponding to 

11.9 A simulated beam current. As can be 

seen from the figure, the difference between the hottest measured ring and the coolest is about 40 C. Though the 

difference in temperatures between the various rings is relatively small, the temperature profile is somewhat 

different from the observed current profiles discussed earlier. In this case, the temperature peak is at ring 5, not ring 

3 as was observed from the current profiles. In general, power deposition should follow the current density 

variations (see Figure 10b). The temperature profile does not follow trends in the current density. The difference 

between the two profiles can be reconciled if one considers the geometry of the magnet rings. Rings 3 and 4 are

mounted on the conical section and are thus canted such that the view factors intersect the optics, which allows for 

radiative cooling as the radiation can actually escape. Ring 5 on the other hand is mounted on the cylindrical section 

and thus the ring’s view factor includes its opposite side and hot discharge chamber wall material. In this regard, 

radiation losses through the grid is not direct, thus it runs hotter. The temperature difference between ring three and

ring four is likely due to differences in current density. All in all, the thermal data reflects the notion that current 

collection alone is not the only driver in operating temperature, rather geometrical considerations are also important 

as apparently demonstrated in this work. 

V. Concluding Remarks

The discharge of a 50 cm ion thruster was characterized using simulated beam extraction. The discharge 

chamber is designed for single engine operation up to 25 kW. Discharge losses as low as 110 W/A were observed. 

Low discharge losses are desirable at intermediate powers in that it provides a pathway to higher thrust to power. It 

was observed that the current collection profile at the various magnet rings did not vary appreciably over a wide 

flow rate and discharge power range. Magnet temperatures remained well below stabilization limits as the discharge 

Figure 12. Temperature profiles at 11.9 A simulated beam

NASA/CR—2013-217828 11
was operated up to 11.9 A of simulated beam extraction. Flatness profiles were generally good, ranging from 0.7 to 

0.8, with the lower end corresponding to higher utilization. Plasma density in the cylindrical section was high with 

density values approaching 1012/cc. Future work will involve validation of simulated conditions at power levels 

ultimately up to 20 kW.

References

1

Reed, K. and Willenberg, H., “Early commercial demonstration of space solar power using ultra-lightweight arrays,” Acta 

Astronautica, 83, 1250, 2009.

2 N. Wyrsch, D. Dominé, F. Freitas, L. Feitknectht, J. Bailat, C. Ballif, G. Poe, K. Bates, K. Reed, Ultra-light amorphous silicon 

cell for space applications, in: Fourth World Conference on Photovoltaic Energy Conversion, Waikoloa, HI, May 7–12, 

2006.

3

“Boeing moving FAST on Satellite Power and Propulsion System,” Defense Industry Daily, June 14, 2009.

4

Van Noord, J. L., “ Lifetime Assessment of the NEXT Ion Thruster,” Proceedings of the 43rd Joint Propulsion Conference, 8-11 

July 2007, AIAA Paper 2007-5274, 2007.

5 Polk, J.E. et al, “A high power ion thruster for deep space missions,” Rev. Sci. Instr. 83, 073306, 2012.

6

Patterson, M.J., Pinero, L., and Sovey, J.S., “Near-Term High Power Ion Propulsion Options for Earth-Orbital Applications.” 

Proceedings of the 45th Joint Propulsion Conference, AIAA Paper 2009,-4819, 2009.

7

Patterson, M. J. and Rawlin, V.K., “Performance of 10-kW class xenon ion thrusters,” AIAA Paper 88-2914.

8

Goebel, D., “ Ion source discharge performance and stability,” Phys. Fluids, 25, 1093, 1982.

9

Patterson, M.J., Roman, R.F., and Foster, J.E., “Ion Engine Development for Interstellar Precursor Missions,” AIAA 2000-3811.

10 Wirz, R. and Goebel, D., “ Ion Thruster Discharge Performance per Magnetic Field Topography,” Proceedings of the 42nd

AIAA Joint Propulsion Conference, AIAA Paper 2006-4487, 2006.

11Arakawa, Y., “Reduction of plasma loss to discharge chamber walls in a ring-cusp ion thruster,” Journal of Propulsion and 

Power, Vol. 3, 90, 1987.

12Brophy, J.R., “ Simulated ion thruster operation without beam extraction,” AIAA 90-2655, 21st International Electric 

Propulsion Conference, Orlando, FL, 1990.

13Goebel, D.M., Polk, J., and Sengupta, A., “Discharge Chamber Performance of the NEXIS Ion Thruster,” Proceedings of the 

40th AIAA Joint Propulsion Conference and Exhibit, AIAA Paper 2004-3813, 2004.

14Herman, D., Soulas, G.C., and Patterson, M.J., “Status of the NEXT Ion Thruster Long Duration Test After 10,100 Hr and 207 

kg Demonstrated,” NASA TM 2008-215030.

NASA/CR—2013-217828 12

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4. TITLE AND SUBTITLE

Simulated Beam Extraction Performance Characterization of a 50-cm Ion Thruster Discharge 

5a. CONTRACT NUMBER

NNX10CF59P; NNX11CC56C 

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6. AUTHOR(S)

Foster, John, E.; Hubble, Aimee; Nowak-Gucker, Sarah; Davis, Chris; Peterson, Peter; Viges, 

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14. ABSTRACT

A 50 cm ion thruster is being developed to operate at >65 percent total efficiency at 11 kW, 2700 s Isp and over 25 kW, 4500 s Isp at a total 

efficiency of >75 percent. The engine is being developed to address the need for a multimode system that can provide a range of thrust-topower to service national and commercial near-earth onboard propulsion needs such as station-keeping and orbit transfer. Operating 

characteristics of the 50 cm ion thruster were measured under simulated beam extraction. The discharge current distribution at the various 

magnet rings was measured over a range of operating conditions. The relationship between the anode current distribution and the resulting 

plasma uniformity and ion flux measured at the thruster exit plane is discussed. The thermal envelope will also be investigated through the 

monitoring of magnet temperatures over the range of discharge powers investigated. Discharge losses as a function of propellant utilization 

was also characterized at multiple simulated beam currents. Bulk plasma conditions such as electron temperature and electron density near 

engine centerline was measured over a range of operating conditions using an internal Langmuir probe. Sensitivity of discharge performance 

to chamber length is also discussed. This data acquired from this discharge study will be used in the refinement of a throttle table in 

anticipation for eventual beam extraction testing. 

15. SUBJECT TERMS

Propulsion; Electric propulsion; Advanced propulsion; In-space propulsion; Thruster; Xenon; Ion propulsion 

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