Jonathan L. Van Noord

Glenn Research Center, Cleveland, Ohio

Lifetime Assessment of the NEXT Ion Thruster

NASA/TM—2010-216915

November 2010

AIAA–2007–5274
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Jonathan L. Van Noord

Glenn Research Center, Cleveland, Ohio

Lifetime Assessment of the NEXT Ion Thruster

NASA/TM—2010-216915

November 2010

AIAA–2007–5274

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

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NASA/TM—2010-216915 1

Lifetime Assessment of the NEXT Ion Thruster 

Jonathan L. Van Noord

National Aeronautics and Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Abstract

Ion thrusters are low thrust, high specific impulse devices with required operational lifetimes on the 

order of 10,000 to 100,000 hr. The NEXT ion thruster is the latest generation of ion thrusters under 

development. The NEXT ion thruster currently has a qualification level propellant throughput 

requirement of 450 kg of xenon, which corresponds to roughly 22,000 hr of operation at the highest 

throttling point. Currently, a NEXT engineering model ion thruster with prototype model ion optics is 

undergoing a long duration test to determine wear characteristics and establish propellant throughput

capability. The NEXT thruster includes many improvements over previous generations of ion thrusters, 

but two of its component improvements have a larger effect on thruster lifetime. These include the ion 

optics with tighter tolerances, a masked region and better gap control, and the discharge cathode keeper 

material change to graphite. Data from the NEXT 2000 hr wear test, the NEXT long duration test, and 

further analysis is used to determine the expected lifetime of the NEXT ion thruster. This paper will 

review the predictions for all of the anticipated failure mechanisms. The mechanisms will include wear of 

the ion optics and cathode’s orifice plate and keeper from the plasma, depletion of low work function 

material in each cathode’s insert, and spalling of material in the discharge chamber leading to arcing.

Based on the analysis of the NEXT ion thruster, the first failure mode for operation above a specific 

impulse of 2000 sec is expected to be the structural failure of the ion optics at 750 kg of propellant

throughput, 1.7 times the qualification requirement. An assessment based on mission analyses for

operation below a specific impulse of 2000 sec indicates that the NEXT thruster is capable of double the 

propellant throughput required by these missions. 

Nomenclature

AB beamlet cross sectional area

C general constant

e elementary charge

F thrust

IB beam current

Icx charge exchange ion production current

Ie discharge current

Isp specific impulse

I

++/I+ ratio of doubly to singly charged ions

jb, peak peak beam current density

JNK neutralizer keeper current

ln neutralization length

m total mass flow rate

mdc in  mass flow rate into discharge chamber

no neutral density

n+ ion density
NASA/TM—2010-216915 2

Pin-PPU power into power processing unit

R ratio of net to total accelerating voltage

Va accelerator grid voltage

Vbps beam power supply voltage

vi ion velocity

VNET beam net accelerating voltage

vo neutral velocity

Vol volume of the charge exchange production region

Y sputter yield

β constant based on ion mass

γ sputter rate

σcx ion charge exchange cross section

φsg screen grid ion transparency

Introduction

NASA’s Evolutionary Xenon Thruster (NEXT) project encompasses the development of a high 

performance, nominal 7-kW, Electric Propulsion Thruster (EPT); a light weight, high efficiency power 

processing unit (PPU); a highly flexible advanced Propellant Management System (PMS); and a lightweight, low-cost gimbal (Refs. 1 and 2). The goal of the NEXT project is to develop the next generation 

ion propulsion technology to NASA Technology Readiness Level (TRL) 6 (Ref. 3). The NEXT project 

has built on knowledge gained from the current state-of-the-art NASA Solar Electric Propulsion 

Technology Application Readiness (NSTAR) thruster system to create a more efficient, higher specific 

impulse, and lower specific mass system (Ref. 2). The design approach will provide future NASA science 

missions with the greatest value in mission performance at a low total development cost.

The NEXT thruster is a 36 cm beam diameter, 2-grid ion optics, xenon ion thruster that uses an input 

power of 0.54-6.9 kW and has a maximum specific impulse of over 4100 sec and a maximum thrust 

greater than 230 mN (Ref. 2). A throttle table containing several of the operating points for the NEXT 

thruster is given in Table 1. Several potential missions have been evaluated using the NEXT thruster (Ref. 

4). These missions included a Saturn mission, Neptune mission, a near Earth asteroid return, comet 

rendezvous, Vesta-Ceres rendezvous, Titan direct lander, and a comet surface sample return. A 300 kg

xenon throughput requirement, which corresponds to a 1.23×107 N-s total impulse, is derived from these 

studies with a corresponding 450 kg qualification level (Ref. 1). This paper will provide an assessment of 

the anticipated lifetime of the NEXT thruster components and demonstrate that the NEXT thruster will 

fulfill the mission propellant throughput requirements. 

Previous extended testing by the NSTAR program examined the wear mechanisms for ion thrusters

(Refs. 5, 6, 7 and 8). Several of these wear mechanisms were minimized and mitigated during the 

NSTAR program, but a set of life-limiting mechanisms was established. Based on these related failure 

modes, an initial life assessment of the NEXT thruster was prepared during Phase I of the NEXT program

(Ref. 9). The relevant failure modes evaluated in this study are shown in Table 2. To date several 

iterations of life models have been utilized based on the most recent data available for the NEXT thruster.

These have included deterministic models for the ion optics (Refs. 10, 11, and 12) and the hollow 

cathodes (Refs. 9 and 13). Modeling of the thruster was also accomplished using a probabilistic failure 

analysis (PFA) process described in Reference 14. To date, PFA modeling uncertainties have led to 

unrealistic results that are not supported by test data. Until these uncertainties are resolved, PFA is of 

limited value. Further work is continuing to reduce these uncertainties, such as measurement of the grid 

gap during extended operation (Refs. 15 and 16).
NASA/TM—2010-216915 3

TABLE 1.—NEXT THROTTLE TABLE: DEMONSTRATED THRUSTER 

AND PPU PERFORMANCE AT BEGINNING OF LIFE

Vbps, 

V

F, 

mN

Isp, 

sec

Pin-PPU, 

kW

Ie, 

A

Va, 

V

I

++/I+ JNK, 

A

m = 5.76 mg/sec;

IB = 3.52 A

1800 236 4190 7.220 18.20 –210 0.045 3.0

1567 221 3910 6.385 18.83 –210 0.045 3.0

1396 208 3690 5.780 19.30 –210 0.045 3.0

1179 192 3395 4.965 19.91 –200 0.045 3.0

m = 5.12 mg/sec; 

IB = 3.10 A

1800 208 4150 6.390 16.48 –210 0.045 3.0

1567 194 3875 5.660 17.03 –210 0.045 3.0

1396 184 3660 5.085 17.44 –210 0.045 3.0

1179 169 3360 4.490 17.98 –200 0.045 3.0

m = 4.46 mg/sec; 

IB = 2.70 A

1800 181 4150 5.600 14.72 –210 0.045 3.0

1567 169 3875 4.920 15.20 –210 0.045 3.0

1396 160 3660 4.455 15.56 –210 0.045 3.0

1179 147 3360 3.860 16.03 –200 0.045 3.0

1021 137 3125 3.425 16.41 –175 0.045 3.0

m = 3.92 mg/sec; 

IB = 2.35 A

1800 158 4100 4.870 13.14 –210 0.045 3.0

1567 147 3830 4.315 13.57 –210 0.045 3.0

1396 139 3615 3.910 13.87 –210 0.045 3.0

1179 128 3325 3.390 14.28 –200 0.045 3.0

1021 119 3090 3.010 14.61 –175 0.045 3.0

m = 3.16 mg/sec;

IB = 2.00 A

1800 134 4310 4.235 13.88 –210 0.063 3.0

1567 125 4025 3.760 14.12 –210 0.063 3.0

1396 118 3800 3.415 14.35 –210 0.063 3.0

1179 108 3490 2.970 14.72 –200 0.063 3.0

1021 101 3250 2.620 15.07 –175 0.063 3.0

m = 2.60 mg/sec;

IB = 1.60 A

1800 107 4190 3.460 12.45 –210 0.063 3.0

1567 99.9 3910 3.080 12.65 –210 0.063 3.0

1396 94.3 3690 2.765 12.83 –210 0.063 3.0

1179 86.7 3395 2.415 13.13 –200 0.063 3.0

1021 80.6 3155 2.160 13.40 –175 0.063 3.0

m = 2.05 mg/sec;

IB = 1.20 A

1800 80.2 4000 2.585 8.36 –210 0.063 3.0

1567 74.9 3735 2.300 8.50 –210 0.063 3.0

1396 70.7 3525 2.090 8.62 –210 0.063 3.0

1179 65.0 3240 1.825 8.83 –200 0.063 3.0

1021 60.4 3015 1.635 9.02 –175 0.063 3.0

936 57.8 2885 1.520 9.15 –150 0.063 3.0

850 55.1 2745 1.415 9.29 –125 0.063 3.0

679 49.2 2450 1.210 9.54 –115 0.063 3.0

650 48.1 2400 1.175 9.54 –144 0.063 3.0

400 37.2 1855 0.865 9.54 –394 0.063 3.0

300 31.8 1585 0.740 9.49 –525 0.063 3.0

m = 1.85 mg/sec; 

IB = 1.00 A

275 25.5 1400 0.610 7.99 –500 0.038 3.0

Several tests of the NEXT thruster and its components to evaluate wear mechanisms have already 

been completed. The NEXT Engineering Model 1 (EM1) thruster was wear-tested for 2038 hr (43 kg of 

xenon throughput) to evaluate thruster wear (Ref. 17). Two wear mechanisms were revealed that could 

potentially lead to a premature failure. One was the excessive wear on the discharge cathode keeper 

orifice plate and the other was excessive wear on the accelerator grid holes beyond a radius of 15.3 cm.

While loss of the keeper orifice plate does not result in thruster failure, it does protect the cathode and 

extend its life. The keeper material was changed to graphite for subsequent designs to achieve significant 

margin for life. The ion optics design was modified by removing some of the outer holes and making the 

effective beam diameter 36 cm instead of the previous 40 cm. The removal of holes was only in regions 

of low plasma density resulting in effectively similar discharge losses without the added optics wear.
NASA/TM—2010-216915 4

TABLE 2.—SUMMARY OF CREDIBLE ION THRUSTER FAILURE MODES

Component Wear mechanism

Discharge cathode Insert barium depletion resulting in inability to ignite or excessive temperatures

Wear of orifice plate causing structural failure

Wear of keeper orifice plate resulting in exposure of cathode orifice plate and heater 

Neutralizer cathode Insert barium depletion resulting in inability to ignite or excessive temperatures

Excessive wear of orifice plate 

Wear of keeper tube exposing cathode tube and heater

Neutralizer orifice clogging preventing proper cathode operation

Cathode heaters Mechanical failure from cyclic operation

Accelerator grid Pit and groove wear from charge exchange ions that result in structural failure

Aperture enlargement resulting in electron backstreaming

Screen grid Upstream wear from impinging ions

Ion optics assembly An unclearable short between grids from debris present during launch or spalled material from wear

Discharge chamber Poor flake retention of sputtered material resulting in optics arcing

Magnet temperatures excessive resulting in field degradation

Insulators and HVPI’s Degrading voltage standoff from extended operation at elevated temperature

Currently, the NEXT Engineering Model 3 (EM3) is undergoing a long duration test (LDT) at the 

NASA Glenn Research Center (Refs. 15 and 16) The intent of this test is to validate and qualify the 

NEXT propellant throughput capability at the highest power throttling point to a qualification-level of 

450 kg. The operation to date has been at a beam power supply voltage of 1800 V and a beam current of 

3.52 A. Currently, EM3 has processed 207 kg of xenon and operated for 10,100 hr. As the data from this 

test continues to become available, it will be incorporated into the life assessment. 

The purpose of this document is to provide a summary of the life assessment for the NEXT ion 

thruster and its various components. This assessment will continue to evolve with time based on the latest 

test data and modeling, but this study represents the best understanding of the anticipated life of the 

thruster and its components. Table 2 contains a list of credible ion thruster failure mechanisms that will be 

examined in this paper. A review of previous ion thruster endurance tests along with the causes of failure 

is given in Reference 18.

Service Life Assessment of the NEXT Thruster

NEXT Discharge Cathode Lifetime 

The discharge cathode provides the electrons to ionize the propellant gas. The discharge cathode is a 

refractory metal tube containing an impregnated porous tungsten insert with a low work function material 

for electron emission at reduced operating temperatures. It is commonly referred to as a hollow cathode.

A swaged heater is used initially for discharge ignition, but once ignition occurs, the insert is self-heated 

by the plasma. The hollow cathode is enclosed in a keeper that protects the cathode from the plasma 

contained in the discharge chamber. The discharge cathode has multiple mechanisms that affect its 

lifetime. One mechanism is the depletion of the insert impregnated material resulting in an increased work 

function. This would lead to significantly higher insert temperatures, higher discharge voltages, and 

potential difficulty in ignition. Another lifetime mechanism is excessive wear of the cathode and keeper 

orifice plates from impinging ions. Significant wear of the cathode orifice plate would be preceded by the 

keeper orifice plate eroding sufficiently to expose the cathode orifice plate. The cathode orifice plate 

could then be eroded to the point of compromising welds, losing the orifice plate, and sputter erosion of 

the sheathed heater causing it to fail and prevent further ignitions. These lifetime mechanisms will be 

evaluated in the following sections.
NASA/TM—2010-216915 5

Insert Barium Depletion

There are two primary factors that determine the life of an impregnated insert in the hollow cathode.

They include the formation of tungstates, which could bind up the barium impregnate and inhibit 

diffusion, and depletion of barium impregnate from the insert (Ref. 19). The current hollow cathode life 

models are either based on predictions of chemical processes and vapor pressures (Refs. 13 and 20), or on 

lifetime data from planar impregnated cathodes operated in a vacuum tube (Ref. 21). The chemistry-based 

model is limited due to the shear complexity of the chemical processes at the insert and the limited high 

temperature chemistry data in the model. There are lifetime data compiled for planar impregnated 

cathodes used in vacuum tubes (Ref. 22), however, due to the different operating environments and 

mechanical designs (cylindrically enclosed versus planar) it is unclear how well the planar vacuum data

and life models transport to a hollow cathode. In all of these models, however, emitter temperature is the 

major factor for determining the insert’s lifetime (Refs. 19 and 23) The approach here will be to use the 

planar impregnated cathode model relationship between insert temperature and lifetime while confirming 

its results with previous cathode life tests.

This model is based significantly on temperature and life data of the planar impregnated cathodes

derived from a relationship outlined by Palluel and Shroff (Ref. 23). These cathodes were operated in a 

vacuum environment. The model examined the relationship between temperature, barium depletion depth 

and lifetime for impregnated planar cathodes that are used in the vacuum industry. These relationships are 

then used to estimate hollow cathode emitter lifetime based on temperature and insert thickness (Ref. 21).

While it appears possible that barium removal may not be significantly affected by the plasma 

environment of the hollow cathode compared to the vacuum environment (Ref. 24), it is still uncertain to 

what degree the negative impact on life would be from the tungstates that form (Ref. 19), or the positive 

impact on life that results from the reuse of barium by effectively enclosing the insert in the cathode tube

(Ref. 20).

To establish the credibility of any model, it is important to compare the results to known data. For a 

hollow cathode insert, it is important to establish a maximum depth that barium can be depleted and yet 

efficiently emit electrons. At the completion of the NSTAR Extended Life Test (ELT), the thruster 

operated for more than 30+ khr and both the neutralizer and discharge cathodes were still operational

(Ref. 5). Measurements were taken to determine barium depletion in the insert to a depth of 500 μm. The 

ELT results showed some barium depletion at depths of 100 and 300 μm at the downstream end of the 

insert in the discharge chamber cathode, but not at 500 μm. So a cathode should operate adequately with 

barium depletion depths of at least 300 μm.

Another hollow cathode was tested at 12 A of emission current for 28 khr before it failed to ignite

(Ref. 25). The peak insert temperature measured on a similar cathode was 1180 °C at 12 A of emission 

current (Ref. 26). Using that peak temperature and an estimated depletion depth of 500 μm, the model 

outlined in References 21 and 23 yields a lifetime estimate of 30 khr. While it is difficult to draw further 

conclusions when only one test was compared to the model, it does lend credence to the model and a 

rational for a barium depletion depth of 500 μm resulting in end of insert life, which is consistent with the 

NSTAR ELT results.

The useful life of the NEXT cathode inserts (both discharge and neutralizer) was calculated. The 

inserts were assumed to only have a thickness of 500 μm to extract barium, even though the actual 

thickness is larger. Peak insert temperature is the other parameter needed for the model. Currently, no 

insert temperature measurements are available for the NEXT insert, but orifice plate temperatures have 

been measured on a cathode similar to the discharge cathode and are shown in Figure 1. These 

temperatures were measured using a planar anode configuration. The cathode insert lifetime model uses 

the depletion depth and peak temperature to produce a lifetime estimate in hours for a particular discharge 

current. Based on the NEXT throttle table, the peak cathode emission current would be 20 A. This 

corresponds to an orifice plate temperature of at most 1030 °C, as shown in Figure 1. It has been shown 

that the emitter on a cathode can be up to 90 °C cooler (Ref. 26) or 100 °C hotter (Ref. 27) than the 

orifice plate temperature. For the purposes of this study, the lifetime results are shown over a ± 100 °C 
NASA/TM—2010-216915 6

700

800

900

1000

1100

1200

5 10 15 20 25 30 35

Anode Current, A

Temperature, o

C

Figure 1.—Orifice plate temperatures measured on a cathode similar to 

the NEXT discharge cathode.

Figure 2.—Discharge cathode propellant throughput capability for a given beam current. The numerous data points at 

given beam current are due to different discharge currents required for different beam voltages.

temperature span on the orifice plate temperatures. Using the NEXT throttle table as shown in Table 1, 

the discharge current can be related to the beam current, and the cathode lifetime in hours can be

converted to a thruster propellant throughput. Using results from the planar cathode life data, the

prediction of thruster cathode emitter lifetime in kilograms of xenon throughput for a particular beam 

current is shown in Figure 2. The least amount of propellant throughput for the NEXT thruster is at a 

3.52 A beam current and results in an insert lifetime of at least 1000 kg xenon throughput, which is over 

two times the qualification requirement. This would correspond to an insert temperature around 1130 °C 

and the lifetime of 48 khr. As the beam current decreases, the expected lifetime of the insert will increase.

Keeper Wear

The main purpose of the cathode keeper is to protect the cathode from excessive erosion due to ion 

impingement from the discharge plasma. Early in the NSTAR program, the thruster was wear-tested 

without a keeper and severe erosion of the cathode orifice plate and heater occurred (Ref. 7). Subsequent 

to that test, a sacrificial keeper was added to protect the cathode (Ref. 8). Once the keeper is sufficiently 

eroded to constitute its end of life, the cathode and heater are exposed and their erosion rates dramatically 

increase, but the keeper end of life would not result in the end of life of the thruster. In this study, keeper 

erosion to the point of exposing the cathode and heater will be conservatively considered a failure.

Previously, the NEXT discharge keeper service life had been analyzed based on the NSTAR ELT test 

findings (Ref. 28). This model and analysis were completed with some component level testing and prior 

1.E+03

1.E+04

1.E+05

1.E+06

1.E+07

1 1.5 2 2.5 3 3.5 4

Beam Current (A)

Propellant Throughput (kg) 

Peak Insert Temp.=Orifice Plate Temp. -100 °C

Peak Insert Temp.=Orifice Plate Temp.

Peak Insert Temp.=Orifice Plate Temp.+100 °C
NASA/TM—2010-216915 7

to any wear test of a NEXT thruster. The assessment determined that the keeper should last a minimum of 

375 kg of xenon throughput and possibly up to 844 kg. However, during the NEXT 2000 hr wear test, 

higher than anticipated keeper wear was found (Ref. 17). The cross section of the keeper after the test is 

shown in 

Figure 3. The wear profile can be seen and the maximum thickness eroded is roughly at a radius of 40

percent of the total keeper radius. This profile is similar to those of previous thruster tests (Ref. 6). The 

maximum keeper erosion from the NEXT 2000 hr test was measured to be about 15 percent of the keeper

orifice plate thickness. With the conservative assumption that the eroded depth exhibited in that valley 

would be linear with time, the keeper orifice plate propellant throughput capability was estimated to be 

280 kg. While this would not constitute a thruster failure, it allows for excessive erosion of the cathode 

orifice plate and heater to occur sooner than desired. To ensure ample lifetime, the discharge keeper 

material was changed to graphite, which has a sputter yield that is about 15 percent of the original 

material.

To predict the discharge keeper lifetime, a new keeper erosion model has been developed. Currently, 

the model is based on erosion profiles from wear tests and attempts to evaluate the angular impact of ions 

on the keeper face. The primary assumption of this model is that the ions wearing the keeper originate

from an effective point source, which is depicted in Figure 4. This assumption is not fully validated by 

experimentation, but the model does demonstrate that the wear pattern could be caused by angular 

dependence of the impinging ion. The model assumes that the ion current density on the keeper drops by 

the square of the distance from this point source and the total current to the keeper must equal the 

measured ion current. The ion current also is assumed to have a fixed ratio of doubly and singly charged 

ions. Another assumption used is the angular dependence of wear based on the angle of impact from a 

xenon ion originating from the point source onto the keeper surface. Currently, the angular dependence is 

based on the lowest measured energy profile available of 78 eV xenon ions and is shown in Figure 5. This

profile is then scaled down to provide the sputter yield as a function of impact angle for lower ion 

energies. Xenon ions impact the keeper at energies around 25 and 50 eV for singly and doubly charged 

ions, respectively. Ions with energies of 25 and 50 eV have normal sputtering yields approximately 0.5 to 

50 percent of the 78 eV ions, respectively (Ref. 29). This model cannot determine the relative amounts of 

singly or doubly charged ions causing the wear, so only a single bulk sputter yield will be used to 

represent their net effect. The model does not incorporate additional physical parameters such as electric 

and magnetic fields, so it cannot predict the point source location a priori. This requires a fit to 

experimental wear data in order to predict the wear rate with time. 

Figure 3.—NEXT keeper orifice cross-section. Picture aspect ratio is changed 

to exaggerate wear profile.

Figure 4.—Picture depicting ion point source approximation for keeper 

wear model.

Point source 

approximation Ion trajectories

Keeper face

Angle of 

impact

Keeper 

orifice

Approximate original 

keeper profile
NASA/TM—2010-216915 8

Figure 5.—Data showing the angular dependence of sputter yield for 

78 eV xenon ions impacting on molybdenum (Ref. 30).

Figure 6.—Comparison of keeper wear model predictions to NSTAR 

8200 hr test keeper cross-section.

The results of the model for the NSTAR keeper are shown in Figure 6. The keeper cross section

shown is from the NSTAR 8200 hr test (Ref. 6). It is assumed that 0.2 A of ion current is collected 

by the keeper (Ref. 31). The sputter yield values used to produce Figure 6 had a normal value of 

2.3×10–5 atoms/ion, which is 0.27 percent the values that are shown in Figure 5. If there were only singly 

charged xenon ions, this normal sputter yield would correspond to ions with an energy around 20 eV 

(Ref. 29). This calculated ion energy corresponds roughly to a plasma potential of 25 eV (assuming the 

keeper is around 5 V of cathode potential) and demonstrates feasibility of the model since plasma 

potentials have been measured near the cathodes to a range of 10 to 30 eV (Ref. 32).

The predicted keeper downstream profiles shown in Figure 6 indicate a reasonable correlation and

that the angular dependence of the impacting xenon ion on the keeper could be a significant factor when 

estimating a keeper’s wear and lifetime. Additional work is needed to improve the predictions on the 

inner diameter of the orifice region. It is possible that the point source will retreat towards the keeper once 

it begins eroding and increase the erosion closer to the orifice. Likely, the orifice will erode open faster

than predicted after ~ 16 khr once a conical shape is established in the entire orifice region. A majority of 

the keeper is predicted to be worn away after 28 khr of operation as shown in Figure 6. Unfortunately, the

keeper orifice wear exhibited during the NSTAR ELT does not offer useful data to compare to the 

model’s wear prediction since a cathode to keeper electrical short occurred early during the test that

resulted in accelerated erosion (Ref. 33). It is valuable to note that even with the increased erosion during 

the test penetration through the face of the keeper was never observed.

0

0.01

0.02

0.03

0.04

0.05

0.06

0.07

0.08

0 10 20 30 40 50 60 70 80

Ion Impact Angle, degree

Sputter Yield, atoms/ion

0.0

0.2

0.4

0.6

0.8

1.0

0.25 0.35 0.45 0.55 0.65 0.75 0.85 0.95

Radial Position

Height

12 khrs

4 khrs

16 khrs

8 khrs

20 khrs

24 khrs 28 khrs 30 khrs

Initial Profile

Profile used in 

model

Experimental data
NASA/TM—2010-216915 9

The model was then used to predict the wear on the NEXT molybdenum keeper. The model 

comparison to the NEXT 2000 hr wear results is shown in Figure 7. Based on measurements, the keeper 

ion current was 0.225 A during that test (Ref. 31). Excellent agreement is seen on the keeper face profile

with the profilometry data taken after the 2000 hr wear test. It is expected that with additional time, the 

region inside the 0.5 radius location will erode faster than the model predicts. Then, just as with the 

NSTAR keeper, the erosion through the entire keeper thickness will progress from the orifice out. To fit 

the experimental wear data, the sputter yield values used had a normal value of 3.4×10–5 atoms/ion and 

were 0.4 percent of the values shown in Figure 5. If there were only singly charged xenon ions, the

normal sputter yield could correspond to an ion energy of around 22 eV (Ref. 29). This value again shows 

that the model is predicting feasible results. The prediction is that the keeper would erode to expose the 

heater after approximately 18 khr of operation. This would correspond to 380 kg of xenon throughput and 

agrees with a pretest prediction of 375 kg (Ref. 17). This is less that than the 450 kg throughput 

qualification level, but does not constitute an end of life, rather the point at which the heater would be 

exposed to the plasma and experience increased wear.

An effort is continuing to improve the keeper model and apply it to the NEXT graphite keeper.

Concurrent to the development of the model, the decision was made in the NEXT project to change to 

keeper material to graphite in order to substantially lower the sputtering yield and ensure adequate 

lifetime. The effect of the change to graphite on the volumetric wear can be determined based on the 

equations described in Reference 28. The change in volumetric wear of the keeper is determined by the 

ratio of the effective yield of the materials based on singly and doubly charged ions, the keeper’s mass, 

and its density. The sputtering yield values for molybdenum and carbon can be found in Reference 29.

The resulting volumetric wear rate for the molybdenum keeper is about 6.5 times the rate for the graphite 

keeper.

There are limited low energy angular sputter data for graphite to predict the angular change from 

molybdenum. The model used a similar angular dependence as molybdenum to predict the graphite 

keeper wear, but with a magnitude that is 6.5 times less. The predicted propellant throughput using a 

graphite keeper would be in excess of 117 khr, which is 2400 kg Xe throughput—5.3 times qualification 

requirement for thruster throughput of 450 kg, as shown in Figure 8. This provides significant margin for 

life. One of the few thruster wear tests to utilize a graphite keeper was the Herakles precursor engine 

HiPEP. During its 2000 hr test, minimal wear on the keeper was found. In fact, what appears to be 

machining marks were still visible (Ref. 34). Using the 6.5 times erosion rate difference, the NEXT 

keeper would only have exhibited around 2 percent thickness reduction in the NEXT 2000 hr test if it had 

been graphite. While a change in keeper orifice plate thickness cannot be determined in situ, the NEXT 

LDT graphite keeper orifice has shown no measurable wear as of 207 kg xenon throughput and 10,100 hr 

of operation (Ref. 15).

Figure 7.—Keeper wear predictions for the molybdenum NEXT keeper.

0.0

0.2

0.4

0.6

0.8

1.0

0.15 0.35 0.55 0.75 0.95

Radial Position

Height

2 khrs

4 khrs 6 khrs

8 khrs

10 khrs

12 khrs

14 khrs

18 khrs

16 khrs

20 khrs

x - Profilometry Data (2038 hours)

Initial Profile
NASA/TM—2010-216915 10

Figure 8.—Graphite discharge cathode wear rate prediction assuming same angular 

relation as molybdenum at 3.52 A beam current and 1800 V beam voltage.

Figure 9.—Graph showing the minimum thruster throughput eroding the NEXT 

keeper exposing the heater.

A prediction of the relative keeper wear between throttle points can be determined based on ion 

keeper current and discharge voltage measurements. The propellant throughput calculated here before the 

cathode heater would be exposed at alternate throttle points was based on the initial prediction of the 

throughput at the peak power. This initial prediction was scaled for other throttle points based on the ratio 

of ion keeper current, sputter yield, and discharge flow rate at the two throttle points. The ion keeper 

current was measured for the NEXT thruster over its throttling range and was found to be a strong 

function of the beam current (Ref. 31). The change in keeper lifetime from the discharge voltage was 

included based on the yield data for C2 (the sputter yield of the C2 molecule is greater than the single atom 

at lower energies) (Ref. 29). The results of using these ratios for various throttle points are shown in 

Figure 9. The spread at each beam current is due to the variation in discharge voltage. It can be seen that 

the observed change in discharge voltage would appear to be an insignificant effect on its lifetime. The 

propellant throughput when the keeper would sufficiently erode to expose the heater increases slightly as 

the beam current decreases. Since the throughput on all throttle points predicted is at least 5 times the 

qualification requirement, the analysis and testing from the LDT indicates that risk of failure due to 

keeper wear is effectively mitigated.

Cathode Orifice Plate Wear/Keeper Shorting

The cathode orifice plate experienced very little wear during the NEXT 2000 hr test (Ref. 35). It 

experienced a slight decrease in diameter from sputter material deposition, but the measured decrease was 

0.0

0.2

0.4

0.6

0.8

1.0

0.15 0.35 0.55 0.75 0.95

Radial Position

Height

Initial Profile

13 khrs

26 khrs 39 khrs

52 khrs

65 khrs

78 khrs

91 khrs

117 khrs

104 khrs

130 khrs

2000

2100

2200

2300

2400

2500

2600

2700

2800

2900

3000

1 1.5 2 2.5 3 3.5 4

Beam Current (A)

Propellant Throughput (kgs)
NASA/TM—2010-216915 11

approximately equal to its uncertainty (Ref. 36). As long as the orifice plate is protected by the keeper, 

minimal wear is anticipated. The cathode orifice diameter decreased slightly from deposited materials, but 

is not expected to change much over the life of the thruster. The orifice diameter in the NSTAR ELT 

essentially remained constant over the 30 khr of operation (Ref. 5). As of 207 kg xenon throughput and 

10,100 hr of operation during the NEXT LDT, the orifice and the cathode orifice plate have shown no 

measurable wear (Ref. 15). This indicates that the cathode orifice plate is sufficiently protected by the 

graphite keeper for at least 2000 kg of thruster propellant throughput and should have a lifetime many 

times the qualification requirement.

During the NEXT 2000 hr wear test, material was deposited on the upstream surface of the keeper

(Ref. 35). This also occurred during NSTAR wear testing (Ref. 6). If the material was assumed to be 

deposited at a linear rate, the gap of the NEXT discharge cathode would be bridged after 560 kg of 

thruster throughput. This would likely lead to an electrical short between the cathode and the keeper and, 

therefore, increased keeper wear, but well beyond the qualification requirement of 450 kg throughput. An 

electrical short is not expected to compromise the ignition or operation of the thruster as confirmed from 

the NSTAR ELT (Ref. 33).

NEXT Neutralizer Cathode Lifetime

The neutralizer on the NEXT thruster is a hollow cathode very similar to the discharge cathode, albeit 

1⁄2 the diameter. The purpose of the neutralizer is to produce the electrons necessary to neutralize the ion 

beam, thereby, keeping the spacecraft from charging. The primary lifetime constraints on the neutralizer 

cathode are the depletion of low work function material from the impregnated insert, which would lead to 

increase temperatures and difficulty in ignition, the neutralizer keeper wear, neutralizer orifice clogging

(Ref. 5), and failure of the heater’s center insulator from cyclic operation.

Insert Barium Depletion

The same model used for the discharge cathode can be implemented for the neutralizer cathode. The 

corresponding temperatures used were based on insert temperature measurements from a similarly sized

cathode (Ref. 26). The temperature data was extrapolated using the thermal model described in 

Reference 37 for operation below 6 A since this was below the lowest current level reported with 

temperature data. Using the NEXT throttle table as shown in Table 1, a propellant throughput capability

based on beam current can be established. This throughput projection is shown in Figure 10 using tighter 

temperature margins than the discharge cathode since direct insert temperature measurements are used 

Figure 10.—NEXT Neutralizer insert lifetime as a function of thruster beam current for insert temperatures ± 50 °C of 

measurements for a comparably sized cathode.

1000

10000

100000

1 1.5 2 2.5 3 3.5 4

Beam Current (A)

Propellant Throughput (kg) 

Peak Insert Temp. = Measured

Insert Temp. - 50 °C

Peak Insert Temp. = Measured

Insert Temp.

Peak Insert Temp. = Measured

Insert Temp. + 50 °C
NASA/TM—2010-216915 12

instead of orifice plate temperatures. The neutralizer insert is anticipated to exceed the discharge insert 

lifetime, since it operates at a cooler temperature, and operate with a thruster xenon throughput in excess 

of 1000 kg. For further comparison, the neutralizer cathode has similar geometry to the cathode that was 

wear tested for 28 khr at 12 A (Ref. 25). So, a realistic expectation would be for the NEXT neutralizer 

insert to operate cooler than the 12 A case and provide well in excess of the 28 khr demonstrated by that 

cathode.

Neutralizer Keeper Wear

The wear on the neutralizer keeper face has generally not been a failure mechanism. The NSTAR 

ELT test actually experienced net deposition on the keeper face (Ref. 5). There was no wear measured on 

the NEXT neutralizer keeper face after the 2000 hr wear test and it is not considered to impact the life of 

the thruster (Refs. 35 and 36).

However, the most susceptible part of the neutralizer keeper is the portion of the tube facing the 

thruster beam. The NSTAR neutralizer tube wall was measured to have eroded up to 20 percent of its 

thickness during the 30,352 hr of operation during the ELT (Ref. 5). The neutralizer keeper wall facing 

the beam in the NEXT 2000 hr test experienced wear up to 7.5 percent through its radial thickness

(Refs. 35 and 36). This would indicate that the neutralizer keeper could wear to the point where a small 

hole could form after 27 khr, which is approximately 570 kg of xenon throughput. However, this hole 

would still not constitute a failure. Any heater or cathode tube wear that would subsequently occur could 

lead to a failure. An analysis of neutralizer wear from direct impingement of beam ions and elastically 

scattered beam ions from the background neutrals was also performed (Ref. 38). In that analysis, the 

axisymmetric code PlumeTool was used to evaluate the thruster plume along with CEX2D. That analysis

found that the major component of the erosion on the side of the neutralizer was from direct impingement 

by plume ions. It was also determined that since the PM ion optics have a reduced beam diameter of 36 

cm when compared to the EM ion optics used in the 2000 hr test, the direct impingement of the beam on 

the neutralizer would be removed and wear on the PM neutralizer keeper side should be negligibly small. 

In addition, the neutralizer keeper tube wall thickness was increased by 50 percent following the 

NEXT 2000 hr test. So if the wear is conservatively assumed to be comparable to the 2000 hr test and 

considered linear, it would now take approximately 850 kg of xenon throughput before a hole would 

form.

Neutralizer Orifice Clogging and Orifice Plate Wear

One of the issues experience during the NSTAR ELT test was the formation of debris in the 

neutralizer orifice that led to higher keeper voltages and decreased flow rate margins to plume mode

(Refs. 5 and 33). The clogging of the neutralizer cathode orifice is difficult to assess since it occurred 

during testing, but at the end of the test no clogging material remained for analysis. The neutralizer orifice 

in the NSTAR ELT appeared to be about 50 percent blocked during a portion of the test when it was 

operating at its lowest throttle point (Ref. 5). While it is unknown what partially blocked the orifice, the 

material was removed when the engine was throttled back up to full power and the neutralizer keeper 

voltage and flow rate margin to plume mode returned to values corresponding to nominal operation.

During the post test analyses, no evidence of deposits could be found.

Without knowledge of what the material was in the orifice or how it occurred during the NSTAR 

ELT, it is not possible to predict whether this would occur on the NEXT neutralizer cathode. However, 

the neutralizer orifice in the NEXT 2000 hr test decreased by approximately 1.1 percent from net 

deposition from facility sputtering and cathode orifice plate material (Ref. 17). To date, no neutralizer 

orifice clogging has been witnessed during the NEXT LDT (Ref. 15). The NEXT neutralizer orifice 

diameter is almost twice as large as the NSTAR neutralizer orifice and higher emission currents are 

experienced. Both of these factors should minimize the likelihood of neutralizer clogging. It is expected 

that clogging will not occur prior to any other neutralizer event such as a possible keeper hole forming at 

850 kg of propellant throughput.
NASA/TM—2010-216915 13

The neutralizer orifice plate in the NEXT 2000 hr test experienced no measurable wear through its 

thickness (Ref. 17). This is consistent with the lack of erosion found on the NSTAR neutralizer orifice 

plate thickness during the ELT (Ref. 5). The final orifice wear from the ELT appeared to have the same 

profile and magnitude of a previous thruster tested to 8200 hr. This would indicate that the orifice wear

occurs until the orifice reaches a steady state configuration and then the wear ceases to continue. The 

minimum orifice diameter of the NEXT neutralizer in the LDT has not changed. Based on the testing to 

date, neutralizer orifice plate wear is not expected to impact the life of the NEXT thruster.

Swaged Heater Cyclic Life

The swaged heaters on the discharge and neutralizer cathodes are used during the ignition process to 

raise the temperature of the impregnated insert to facilitate initial electron emission. Once the cathodes 

are ignited, the heaters are turned off and the cathodes self-heat from the generated plasma. This on-off 

cycling of the heater could lead to mechanical failure of the heater center conductor and requires that the 

heaters be validated for sufficient cyclic lifetime operation. Previously, cyclic life testing was performed 

on the neutralizer heater design for the International Space Station (ISS) Plasma Contactor hollow 

cathode assembly (Refs. 39 and 40). These heaters were validated to have a B10 life of approximately 

6700 cycles (this corresponds to the cycles at which 10 percent of the population of heaters would fail).

The NSTAR heaters were the same design and from the same manufacturer as the tested hollow cathode 

assembly heaters, so cyclic testing was not needed to demonstrate that they met the NSTAR requirements.

However, the NEXT discharge heater design is slightly different than the tested hollow cathode assembly

heater and from a different manufacturer, so the plan is to also validate the PM heater cyclic lifetimes 

through similar testing (Ref. 41).

Accelerator Grid

Two grids are employed to extract the ions generated in the discharge chamber of the ion thruster.

The inner or screen grid is at a potential similar to the discharge chamber and protects the accelerator 

grid, which is at a significantly lower potential, by focusing the ions through the accelerator grid holes.

The accelerator grid has a negative potential with respect to the spacecraft environment to prevent

neutralizer electrons from backstreaming into the discharge chamber.

There are two main accelerator grid specific failure modes related to wear mechanisms. The first is

structural failure due to “pit and groove” erosion between the apertures on the downstream side of the 

accelerator grid. This erosion is caused by charge exchange ions produced just downstream of the 

accelerator grid. These ions are focused into the region between two holes to form grooves, as shown in 

Figure 11, and into the region equidistant to three adjacent holes to form pits. For this study, the 

accelerator grid is assumed to structurally fail when the groove penetrates through the accelerator grid 

resulting in the end-of-life of the thruster. This wear is related to a number of factors including the beam 

current density, neutral density, and accelerator voltage. The greatest amount of wear from pit and groove 

erosion is typically at the center of the grid where the beam density is the highest.

Figure 11.—Pit and groove erosion on 

the accelerator grid.

Aperture

Pit

Groove
NASA/TM—2010-216915 14

The second failure mode is electron backstreaming resulting from aperture barrel erosion. This 

erosion takes place on the inside wall of the aperture from charge exchange ions produced in the region 

between the grids. As the apertures increase in diameter from erosion, the magnitude of the potential at 

the center of the hole will eventually drop and fail to repel the electrons from the neutralized beam. This 

results in electrons backstreaming into the discharge chamber with high energy and quickly damaging

components such as the discharge cathode. However, if it is determined that backstreaming will occur 

after a certain throughput, the accelerator power supply can be adjusted to prevent backstreaming and 

mitigate this issue with only a minor impact on the lifetime due to structural failure.

In this study, the NEXT propellant throughput capability corresponding to the accelerator grid 

structural failure will be determined at full power by extrapolating the results from the NEXT 2000 hr

wear test and the LDT and then comparing these results to predictions from optics modeling. Then, based 

on the predicted propellant throughput capability at full power, the throughput capability at other throttle 

points will be determined based on the relative change of the dominant wear mechanisms. This 

throughput prediction across the throttle table will then be used to predict the amount of groove erosion 

expected based on three different mission profiles.

Structural Failure from Downstream Erosion

2000 hr Wear Test and NEXT LDT Results

The nominal NEXT optics design geometry is the same as the nominal NSTAR optics geometry with 

the exception that the NEXT optics has 1.6 times the beam area and an accelerator grid that is 1.5 times 

thicker than the NSTAR accelerator grid (Refs. 42 and 43). The service life of the NEXT accelerator grid

is increased by the addition of material to its thickness. If the wear rate in the pits of the NEXT grids 

during the 2000 hr test is assumed linear, the pits would wear through the grid’s thickness after 17.5 khr, 

or ~365 kg of xenon, which is 85 kg less than the 450 kg throughput qualification requirement (Ref. 17).

However, as previously mentioned, it is the grooves that would constitute a structural failure. A linear 

extrapolation from the NEXT wear test would indicate that the deepest point of the grooves would wear 

through after 35.5 khr or ~750 kg of xenon, which is 1.7 times the qualification requirement (Ref. 17).

Based on the latest in situ measurements during the NEXT LDT, the groove is predicted to penetrate the 

grid’s thickness between 700 to 800 kg of propellant throughput at the peak operating power (Ref. 15).

Previous Computer Models

Several different computer software models, such as CEX3D (Ref. 44), CEX2D (Ref. 44), ffx

(Ref. 11), erode (Ref. 10), and MICHELLE (Ref. 45), have been used to examine NEXT accelerator grid 

erosion. Table 3 shows the pit and groove penetration of the NEXT accelerator grid based on predictions

from the codes and an extrapolation of the NEXT LDT and 2000 hr wear test results. The models predict 

a groove penetration of 590 kg and 770 kg of xenon at the peak throttling point, which is 140 to 320 kg

more than qualification requirement. However, the CEX3D code is known to over-predict groove wear by 

2 times the observed rates (Ref. 44). Taking that into account would yield an expected throughput of 

greater than 1000 kg for groove penetration. It should also be pointed out that some of the largest areas of 

uncertainty with the optics models pertain to grid gap, beamlet currents and redeposition of sputtered 

material into the pits or grooves. Recently, testing has shown the NSTAR operational or “hot” grid gap to 

be less than most previously assumed (Ref. 46). Incorrect assumptions can easily cause the predictions of 

the codes to vary by tens of percent when predicting wear. This highlights the importance of wear data to 

calibrate the models and further thruster testing with diagnostics to reduce the uncertainties of model 

inputs. Based on a combination of testing and computer modeling, the accelerator grid is anticipated to 

reach end of life at 750 kg.
NASA/TM—2010-216915 15

TABLE 3.—PROPELLANT THROUGHPUT PREDICTIONS OF PIT AND

GROOVE PENETRATION AT FULL POWER ON THE NEXT ACCELERATOR GRID

CEX3D ffx NEXT 2000 hr 

test

NEXT LDT

Pit penetration (kg) 245 610 365 -------

Groove penetration (kg) 590 770 750 700-800

Extrapolation of Wear Test and Analytical Results Over the NEXT Throttle Table

Based on the NEXT 2000 hr wear test, the NEXT LDT, and with the computer models presented 

here, the propellant throughput coinciding with groove penetration at a beam current of 3.52 A and beam

voltage of 1800 V would occur approximately at 750 kg. It is possible to use the relative change in 

dominant factors affecting the pit and groove wear rate to predict throughputs at a variety of throttling 

points on the NEXT throttle table. These dominant factors are the production of charge exchange ions and 

the sputter yield based on ion energy. A first order approximation of the wear rate can be approximated

by Equation (1), where C is a constant assuming that the wear rate is proportional to the production of 

charge exchange ions and Y is the sputter yield of the grid material. 

Wear Rate = C ⋅ Icx ⋅Y (1)

The charge exchange ion production, Icx, can be calculated with Equation (2) where no is the neutral 

density, n+ is the ion density, Vol is the volume of the charge exchange production region, σcx is the ion 

charge exchange cross section, and vi is the ion velocity.

cx no n e Vol cx vi I = ⋅ + ⋅ ⋅ ⋅σ ⋅ (2)

The neutral density immediately upstream of the grids can be approximated by using the mass flow into 

the discharge chamber, m dcin, the beam current, IB, the double to single ion ratio, I

++/I+

, and the neutral 

velocity, vo, leaving through the optics as shown in Equation (

3

).

o

I

I

I

I

dc in B

o v

m I

n 



































+ ⋅

+ − ⋅

= +

++

+

++

mass equiv.

1 2

1  (3)

The ion density can be calculated using the peak beam current density, jb, peak, divided by the ion velocity.

i

b

v

j n , peak + = (4)

The volume of the charge exchange production region can be estimated by assuming a constant beamlet 

cross section, AB, multiplied by the neutralization length, ln, and shown in Equation (5) (Ref. 47). β is a 

constant based on the ion mass, VNET is the beam supply voltage plus the coupling voltage, which is 

measured between the neutralizer common to tank ground, and R is the R-ratio for an ion thruster, which 

is VNET divided by Vbps plus the absolute value of the accelerator voltage.

( ) 0.5

,peak

1.5 1.5 1 3 4















β⋅ ⋅ + − = ⋅ = ⋅

− −

b

NET

B n B j

V R R

Vol A l A (5)
NASA/TM—2010-216915 16

Assuming linear wear, the lifetime of the grids can then be simply computed by dividing the thickness of 

the grid by its wear rate. The corresponding propellant throughput of the thruster can then be calculated 

by Equation (6). 

⋅m 

 



 = Wear Rate

Grid Thickness Throughput (6)

To determine the relative change in propellant throughput, Equations (1) to (6) can be combined with the 

assumption that C, β, the neutral velocity and beamlet cross sectional area are constant to give:

( )

( )

( )

( )

( )

( )

( )

( )

1

2

2

1

2

0.5

, peak

1.5 1.5 1

mass equiv.

1

0.5

, peak

1.5 1.5 1

mass equiv.

1

2

3 4

1 2

1

3 4

1 2

1

Throughput

Throughput

m

m

Y

Y

m I V R R j

m I V R R j

NET b cx

I

I

I

I

dc in B

NET b cx

I

I

I

I

dc in B









⋅ ⋅





















⋅ ⋅ + − ⋅ ⋅ σ





































+ ⋅

+ − ⋅





















⋅ ⋅ + − ⋅ ⋅ σ





































+ ⋅

+ − ⋅

=

=

− −

− −

+

++

+

++

+

++

+

++

(7)

The charge exchange cross section is determined based on the ion energy, which is a function of the beam 

and accelerator voltage on an ion thruster. The relation used is reported in Reference 48. It is assumed 

that the sputter yield is determined primarily by the accelerator voltage and the relationship used is 

reported in References 29 and 49.

Using Equation (7), along with test and analysis results for the 3.52 A, 1800 V operating condition, 

propellant throughputs at other throttle points can be predicted for structural failure of the accelerator 

grid. These predicted throughputs and lifetimes for various specific impulses on the throttle table are 

shown in Figure 12. Figure 13 depicts these same throughputs and lifetimes as a function of thrust. All of 

the operating conditions are predicted to have throughputs exceeding the maximum power throttling point

and qualification requirement except at the three lowest specific impulses and thrusts. The lowest two

throttle points are the only ones predicted to have less than the 450 kg propellant throughput requirement 

with 410 and 320 kg throughput predicted. This method of predicting lifetime and propellant throughput 

agrees well with the ffx predictions at various throttling points. The propellant throughput prediction that 

varies the most between the two methods is at the lowest power. The ffx model predicts groove 

penetration at significantly higher throughput of 1040 kg. However, the ffx model does show much larger 

area pits than at the higher power throttle points and these pits penetrate at 280 kg of throughput. This 

would seem to indicate that the 300 to 400 kg propellant throughput predictions for structural failure are 

conservative predictions.

While propellant throughput capability at the lowest two throttle points is below the 450 kg 

qualification requirement, the driving requirement is really based on a particular mission profile. Of the 

different mission scenarios examined, there were three missions that had operation at the lower three

power throttling points (Ref. 4). These included a rendezvous mission to Vesta-Ceres, a comet sample 

return mission, and a Titan direct lander mission. Figure 14 through Figure 16 show the time spent and 

predicted accelerator groove wear at specific throttle points for these missions. In these figures, the time 

spent at a throttle point represents the total accumulated time among all thrusters and the predicted groove 

wear is based on an individual thruster assuming that the duration is equally divided over the number of

thrusters used. The mission assumption for Vesta-Ceres is that two thrusters are used in sequential order 
NASA/TM—2010-216915 17

with the first thruster operating primarily at the high power throttle points to deliver the spacecraft to the 

asteroid Vesta, with the second thruster then used for much of the lower power throttling to Ceres. For 

simplicity, this analysis assumed equal distribution between the two engines. However, if the duration at

37 mN, which has an individual propellant throughput of 600 kg, is assumed to be on a single engine the

erosion through the accelerator grid would be at most 4 percent of its thickness. Operation at that lowest 

throttle point clearly is not a significant driver for that mission.

The mission that had the longest operation at the lowest specific impulses was the comet sample 

return mission. It was assumed that two thrusters were used on that mission. When its entire mission 

throttle histogram is used to predict grid life, accelerator grid groove erosion will be through 33 percent of 

the grid thickness with less than a quarter of that erosion from operation at or below a thrust of 37 mN.

The Titan direct lander mission assumes one thruster. On that mission, groove erosion from operation at 

or below 37 mN accounts for only 4 percent out of the 26 percent erosion through the grid.

It can be seen that the missions examined using the lower throttle points that violate the 450 kg 

propellant throughput requirement would still yield an accelerator grid structural failure propellant

throughput that was at least two times the mission throughput required. An additional consideration is a

recent recommendation that the accelerator voltage be operated with reduced margin at these lower

thrusts and specific impulses because of the improved perveance margins offered by the PM ion optics

(Ref. 50). The resulting effect on lifetime and throughput at those throttle points is shown in Figure 12

and Figure 13 with open symbols and would result in all of the propellant throughput capabilities 

exceeding the 450 kg throughput requirement.

Figure 12.—The throughput and lifetime predicted before structural failure of 

the accelerator grid due to groove penetration for specific impulses at 

various throttling points. The open symbols represent the values based on 

a suggested modification to the throttle table (Ref. 50).

10

100

1000

10000

1000 1500 2000 2500 3000 3500 4000 4500

Specific Impulse, s

Propellant Throughput, kg

Lifetime, khrs

ffx Prediction, kg
NASA/TM—2010-216915 18

Figure 13.—The throughput and lifetime predicted before structural failure of 

the accelerator grid due to groove penetration as a function of thrust at 

various throttling points. The open symbols represent the values based on 

a suggested modification to the throttle table (Ref. 50).

Figure 14.—Relative groove wear through accelerator grid and time at 

throttle point for Vesta-Ceres mission. Wear at groove from the total 

mission is predicted to be 51 percent of the grid thickness.

10

100

1000

10000

0 50 100 150 200 250

Thrust, mN

Propellant Throughput, kg

Lifetime, khrs

ffx Predictions, kg

0

2

4

6

8

10

12

14

16

18

20

236221208192137119101 81 58 55 49 48 37

Thrust (mN)

Time at Throttle Point (khr)- Gray Bar

0%

5%

10%

15%

20%

25%

30%

Grid Thickness Eroded - Black Bar
NASA/TM—2010-216915 19

Figure 15.—Relative groove wear through accelerator grid and time at 

throttle point for comet surface sample return mission. Wear at 

groove from the total mission is predicted to be 33 percent of the grid 

thickness.

Figure 16.—Relative groove wear through accelerator grid and time at 

throttle point for Titan direct lander mission. Wear at groove from the 

total mission is predicted to be 26 percent of the grid thickness.

Barrel Erosion Resulting in Electron Backstreaming

Modeling of the accelerator aperture barrel erosion and resulting electron backstreaming appears even 

more difficult for the ion optics codes. While some have predicted the initial electron backstreaming limit 

reasonably well, they do not agree with the results from wear testing the NEXT thruster for how that limit 

changes with extended operation. The models have generally predicted that after 45 kg of throughput at 

full power, the backstreaming limit would have decreased by 10 V and after 200 kg of propellant 

throughput the backstreaming limit would have decreased by 20 V due to aperture barrel erosion. Both 

the NEXT 2000 hr test after 43 kg of propellant throughput and the NEXT LDT after 207 kg of xenon 

throughput and 10,100 hr of operation have shown no significant change in the electron backstreaming 

limit (Refs. 15 and 17). Due to the inability of the codes to accurately predict the change in electron 

backstreaming limit, the best indicators are the current wear tests. These tests indicate that failure from 

backstreaming electrons due to barrel erosion will not occur prior to any structural failure of the 

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NASA/TM—2010-216915 20

accelerator grid at the peak power throttle point. However, since the models do not accurately predict the 

trend in the electron backstreaming limit, further work is needed for evaluation of throughput capability at 

other throttle points.

Screen Grid Erosion

The failure possible with the screen grid would be excessive erosion from the discharge plasma 

leading to structural failure of that grid. Early ion thrusters exhibited higher rates of erosion on the screen 

grids (Ref. 7), but subsequent operational changes to the NSTAR thruster related to the screen grid 

potential and discharge voltage have significantly reduced the wear of the screen grid. The NSTAR screen 

grid apertures only exhibited slight chamfering on the upstream side during the 8200 hr test (Ref. 6) and 

less than 60 μm of chamfering during the NSTAR ELT (Ref. 51). It is possible to correlate the net erosion 

of the screen grids between the NSTAR thruster and the NEXT thruster. The net erosion on the screen 

grid per mass of propellant throughput, shown in Equation (8), is dependent on the sputter yield, Y, screen 

grid ion transparency, φsg, beam current, IB, and the propellant flow rate, m .

( ) 



 



 = ⋅ − φ ⋅ γ

propellant throughput

sputtered atoms 1 , m

I Y

m

B

sg   (8)

The NEXT thruster and NSTAR thruster both operate with discharge voltages between 23 to 28 V, which 

would yield ions with similar energies, and sputtering yields, impacting the screen grid (Refs. 15 and 33)

The screen grid ion transparency on the NSTAR optics is between 0.80 to 0.88 and on the NEXT optics is 

0.78 to 0.90 (Refs. 6, 15, and 33). The beam current per total propellant mass flow rate is 1.68 to 

2.10 A/mg/s on the NSTAR thruster and is 1.58 to 1.85 A/mg/s on the NEXT thruster. If these values are 

used in Equation (8), the NSTAR thruster bulk screen grid erosion per mass of propellant throughput 

would be 6 to 14 percent higher than the NEXT thruster. This result, along with the lower doubly charged 

ion content and lower peak beam current density of the NEXT thruster, would indicate that the NEXT

screen grid wear rate is anticipated to be less than that for NSTAR. The screen grid in the NEXT 2000 hr

test exhibited no signs of excessive erosion (Ref. 17). Screen grid erosion on the NEXT engine is at such 

a slow rate that is considered insignificant. The propellant throughput capability based on screen grid 

erosion will be well in excess of the first anticipated failure mode at 750 kg of propellant throughput.

Optics Assembly Electrical Shorting

A primary failure mechanism that could be caused by the discharge chamber is the grid shorting. This 

could be caused by debris or sputtered material bridging the gap between the grids. Brief electrical 

shorting of the grids will not result in failure, but failure would result from an unclearable electrical short.

It is important to take a fourfold approach to prevent failure from electrical shorting. This approach would 

be to minimize sputtering, retain as much sputtered material on the discharge chamber walls as possible,

prevent any sputtered material that isn’t retained be to be of a size such that it cannot bridge the optics 

gap, and provide a process to clear any material that does create an electrical short.

Minimize Sputtering

The cathode keeper and the ion optics are the two primary contributors to sputtered material. As 

previously mentioned, the cathode keeper material was changed to graphite to reduce the wear rate, and 

hence, the rate at which it sputters material into the discharge chamber. Upstream screen grid aperture 

erosion is also a source of sputtered material. However as was discussed previously, screen grid 

operational changes have greatly reduced the amount of sputtering that occurs here. Erosion of the 

accelerator grid in the aperture also produces sputtered material that would be deposited on either of the 

grids or the discharge chamber. This would come from either an under-focused or over-focused beamlet 

resulting in either uniform barrel or hexagonally shaped cross-over erosion, respectively. Under-focused
NASA/TM—2010-216915 21

beamlet ions typically occur at the highest beam current densities, which is also typically at the center of 

the optics. Operation with under-focused beamlets is usually avoided. Uniform barrel erosion can also 

occur from charge exchange ions. Risk of an electrical short from the material sputtered as a result of 

under-focused beamlets or charge exchange ions is minimal.

Cross-over erosion generally occurs in areas of low beam current density and relatively high ion 

optics total voltages. This would typically occur at the outer apertures. Over-focused beamlet ions in the 

low density region impinge on the accelerator grid producing off-center, hexagonally shaped erosion 

patterns. Since this erosion is due to direct impingement, once the grid material is removed where the 

beam impinges, the erosion stops. Since it stops, this erosion does not jeopardize the structurally integrity 

of the grid, but the sputtered material could contribute to electrical shorting. One of the locations material 

sputtered from cross-over erosion is deposited is on the inner diameter of the accelerator aperture 

experiencing crossover erosion. This deposited material could spall and create an electrical short.

During the NEXT 2000 hr test, which was at full power, the accelerator apertures beyond a radius of 

15.3 cm exhibited a hexagonal star-shaped erosion pattern indicative of cross-over erosion (Ref. 17). This 

wear pattern can be seen in Figure 17. The wear was found to be the result of over-focused beamlet ions 

in this low plasma density region impinging upon the accelerator grid and significantly smaller-thannominal accelerator aperture diameters in this region (Ref. 17).

To mitigate this wear and minimize the amount of sputtered material, the PM design of NEXT optics 

reduced the beam extraction diameter to 36 cm. The area outside of 36 cm was the region of low plasma 

density, so there was a minimal effect on the performance of the engine, while the erosion mechanism 

was removed. Another improvement on the PM optics has been that aperture diameters were kept to a 

tighter tolerance throughout the beam extraction area. Crossover erosion is being monitored closely on the 

NEXT LDT (Refs. 15 and 52). A photograph of the edge holes taken after 1477 hr of operation during the 

LDT is also shown in Figure 17. There is no longer any apparent star shaped pattern indicative of crossover erosion on accelerator grid. The hexagonally shaped cross-over erosion has been significantly 

reduced.

The only remaining cross-over erosion experienced at full power on the PM optics was anticipated

and is a “notching” on the outermost apertures as seen in Figure 17. This was witnessed after the NEXT 

2000 hr test and previous NSTAR tests (Ref. 8). Modeling of this notching phenomenon was undertaken 

at the NASA Glenn Research Center using an optics modeling program, which is capable of modeling 

asymmetric geometries (Refs. 45 and 53). This modeling was able to demonstrate that this erosion is a 

result of extracting ions from the region where there are no apertures, while hex-shaped outer edge 

erosion is due to crossover erosion from to low beamlet current density and relatively high extraction 

voltage. The modeling also concluded that this erosion stops when sufficient material is eroded and does 

not impact the extraction or structural lifetime capability of the accelerator grid. This notching erosion 

contributes minimal sputtered material to risk electrical shorting. 

Figure 17.—Photographs of the edge holes taken after 2038 hr of operation during the NEXT 2000 hr test and after 

1477 hr of operation during the NEXT LDT Test.

Slight “notching”

No star-shaped 

erosion patterns 

(LDT)

Note the star-shaped 

erosion patterns 

NEXT 2000 hr Test NEXT LDT
NASA/TM—2010-216915 22

Sputtered Material Retention in Discharge Chamber/Minimize Debris Size

As previously mentioned, there are two primary contributors to sputtered material on the discharge 

chamber walls. These are from the cathode keeper and the ion optics, primarily the screen grid. It is 

predicted that a maximum of 7 μm of carbon from the discharge cathode keeper could coat the discharge 

chamber surface. There has not be a comprehensive study on the adherence of carbon to the wire mesh, 

but the 2000 hr HiPEP engine, which used a graphite keeper and grit-blasted mesh, showed no evidence 

of spalling (Ref. 34).

It is anticipated that the largest contributor during flight to material deposition in the discharge

chamber will be from the ion optics, primarily the screen grid. As demonstrated in the previous section, 

the bulk erosion of the NEXT screen grid per propellant throughput would be less than that experienced 

on the NSTAR thruster. The ratio surface area of the discharge chamber to the screen grid area is roughly 

50 percent larger on the NSTAR thruster. These two parameters offset each other when determining a 

crude prediction of the relative film thickness expected on the discharge chamber walls. So, the NSTAR 

wear tests provide deposition rates to a first order for the NEXT thruster discharge chamber. Based on the

same deposition rates from the NSTAR wear tests, it is predicted that the maximum deposition rate in the 

discharge chamber would be 7 μm per 100 kg of thruster xenon throughput. Testing has demonstrated that 

up to 130 μm of molybdenum can be sputter deposited onto the grit-blasted wire mesh of the NEXT

discharge chamber without spalling off (Ref. 54). This would mean at least 1850 kg of xenon would need 

to be processed before reaching 130 μm of material deposited on the discharge chamber.

If any material would spall off the discharge chamber walls prior to or after a film thickness of 

130 μm, the wire mesh used on the surface of the discharge chamber is designed to prevent any material 

spalling of sufficient size to electrically short the optics (Ref. 55). The wire mesh is designed with wire 

diameters and gaps small enough so that anything spalling should have sizes of that magnitude. This 

should result in any spalling to be of insufficient size to short between the grids and is a consistent design 

with the NSTAR and NEXT ion thrusters.

Clearing Debris in Optics

Even after minimizing the amount of thruster generated material, it is still possible to get a sustained 

electrical short on the optics. This could be caused by contaminants from spacecraft assembly or launch 

vehicle debris as well as from the thruster. For example, the NSTAR ion thruster on the Deep Space One 

mission is thought to have suffered from this type of contamination (Ref. 56). In that mission, the thruster 

operated at the minimum throttling point for 4.5 min before the thruster experienced continuous recycling 

and shut down. Later the thruster ignited properly without continuous recycling and the short appeared to 

have been cleared by thermal cycling of the thruster.

Since there are various possible sources causing electrical shorting of the grids it is necessary to have 

a grid debris clearing capability. The NEXT approach to clear the debris is similar to that of the NSTAR 

thruster of the DS1 spacecraft and is accomplished by connecting the discharge power supply across the 

screen and accelerator grids to provide the necessary clearing power (Ref. 57). An investigation was 

undertaken to evaluate the currents and voltages required to vaporize materials commonly produced by an 

ion thruster (Ref. 58). Alternative methods to clear sample wires from the grids were also investigated, 

including the use of a capacitor grid clear circuit. It was recommended to initially attempt to clear debris 

using mechanical means such as thermally cycling the thruster. If that fails, it was determined that the 

discharge power supply capability is powerful enough to clear the flakes found in the NEXT discharge 

chamber after the 2000 hr wear test.

It is not possible to provide a propellant throughput prediction related to optics shorting since an 

electrical short could be caused by spacecraft or launch vehicle debris generated during launch. The risk 

of failure from an optics electrical short is minimized by reducing debris generation and adopting grid 

clearing capability and procedures that have been developed on previous ion thruster missions. The 

NEXT program has done this and should ensure operation well beyond 450 kg of propellant throughput.
NASA/TM—2010-216915 23

Discharge Chamber Magnets

Another issue with previous thrusters was the thermal risk to the permanent magnets. At elevated 

temperatures permanent magnets can degrade and decrease in magnetic strength resulting in thruster 

performance degradation. The pre-test and post-test magnetic field mapping of the NEXT discharge 

chamber for the 2000 hr test showed no such degradation (Ref. 17). There is also no change in discharge 

performance of the NEXT thruster during the LDT that would indicate a degradation of the magnetic 

circuit (Ref. 15). Additionally, the NEXT PM thruster operating with the maximum internal heating and

under the hottest mission scenario is predicted to have maximum magnet temperatures of 250 to 288 °C, 

which is slightly cooler than the maximum operational temperatures of 291 °C on the magnets of the 

NSTAR ion engine (Refs. 59 and 60). In addition, the NEXT magnets were stabilized to a higher 

temperature than the magnets used on the NSTAR ion engine. The higher stabilization temperatures and 

slightly cooler operating temperatures assure that the NEXT magnets will degrade slower than the 

NSTAR magnets, which have demonstrated 30+ khr of operation without degradation (Ref. 5). The 

magnets on the NEXT thruster should demonstrate adequate magnetic strength beyond the first thruster 

failure mode at 750 kg of propellant throughput.

High Voltage Propellant Isolators

The high voltage propellant isolators (HVPI) are used in the xenon feed lines to the discharge 

chamber. The risk with HVPI is that at excessive temperature and voltage, a leakage current across the 

isolator would increase to unacceptable levels with extended use. This increased current may not 

constitute a thruster failure, but an increase in power loss in the thruster system. The NEXT HVPI design 

concept is based on the NSTAR HVPI that was successfully tested at the component level and thruster 

level for 16,000 and 30,000 hr, respectively (Refs. 61 and 33). This HVPI design philosophy employed on 

the NEXT thruster is also the same as the NSTAR thruster, but procured from a different manufacturer

(Refs. 62 and 63). The risk of failure is low if the manufacturing processes and materials similar to those 

of the NSTAR isolators. Design verification tests are currently being conducted on the NEXT HVPI

(Ref. 41). At the date of this writing, the HVPI design verification test has operated for more than

10,300 hr at 260 °C, 2300 V, and with an internal pressure of 2.7 kPa and have met the expected 

performance characteristics. The HVPI’s are currently showing a trend of 0.1 nA per 1600 hr (Ref. 41). A

leakage current maximum of 100 μA is the requirement based on NSTAR thruster (Ref. 41). Based on the 

current trends of the test, that definition of failure would yield a propellant throughput significantly 

beyond the first thruster failure mode anticipated for the thruster at 750 kg propellant throughput.

Thruster Summary

A summary of the current lifetime assessment for the NEXT thruster components is shown in Table 4

for operation at the peak power throttle point, which typically corresponds to the least capable propellant 

throughput. A reference to the figures showing how the propellant throughput capability scales across the 

throttle table for the mechanisms is included in the last column. As testing continues, the models will be 

increased in fidelity along with this assessment.
NASA/TM—2010-216915 24

TABLE 4.—TABLE CONTAINING A SUMMARY OF THE NEXT LIFETIME ASSESSMENT

Component Failure mechanism Minimum xenon throughput, kg

(x 450 kg qualification)

Discharge cathode Insert barium depletion 1450 kg (3.2 x), Figure 2

Orifice plate wear >2000 kg (4.4 x)

Keeper wear >2000 kg (4.4 x), Figure 9

Neutralizer cathode Insert barium depletion 1600 kg (3.6 x), Figure 10

Orifice plate wear > 850 kg (1.9 x)

Keeper side wear penetration 850 kg (1.9 x)

Cathode orifice clogging > 850 kg (1.9 x)

Accelerator optics Pit and groove wear / structural failure 750 kg (1.7 x)*, Figure 12

Aperture enlargement / electron backstreaming > 750 kg (1.7 x) †

Screen optics Excessive wear >> 750 kg (1.7 x)

Optics assembly An unclearable short Mitigation through retention of sputtered 

material, thermal cycling of optics, and 

power supply for clearing.

Discharge chamber Flake retention of sputtered material 1850 kg (4.0 x)

Magnet temperatures > 750 kg(1.7 x)

High voltage 

propellant isolators

Degrading voltage standoff from extended operation 

at elevated temperature

>> 750 kg (1.7 x);

Design verification test ongoing.

*This does not include the lowest three power throttle points which are shown to have more than adequate life for anticipated missions. †

As indicated for 3.52 A beam current and 1800 V beam voltage. Other operating points need further investigation.

The first failure mode predicted on the thruster would be accelerator grid structural failure from pit 

and groove erosion. The corresponding propellant throughput at full power would be 750 kg at specific 

impulses above 2000 sec, which well exceeds the 450 kg throughput requirement by 70 percent. This is 

supported by the NEXT 2000 hr wear test erosion measurements and LDT in situ measurements. As the 

specific impulse or thrust decrease, the propellant throughput increases until a specific impulse of 

1854 sec and a thrust of 37 mN. The propellant throughput before structural failure corresponding to the 

two lowest power throttle points on the throttle table is below the requirement of 450 kg. However, these

throttle points are used only on a small number of missions studied. Based on an analysis of the 

accelerator grids for the three missions using the three lowest power throttle points, the total depth of 

groove penetration from the entire mission with the greatest wear is anticipated to be at most 51 percent

of the accelerator grid thickness. This would yield the anticipated lifetime of the accelerator grid before 

structural failure to be about 2 times the mission requirement. There is a degree of uncertainty with the 

2–D and 3–D models due to the lack of operational grid gap measurements, beamlet current densities, 

sputter yield, and redeposition of sputtered material. However, based on erosion measurements from the 

tests, the analysis can be correlated. The models have not been able to accurately predict the change in 

electron backstreaming with time. The models predict noticeable decrease in electron backstreaming 

margin with time at the peak power throttling point, but the tests to date do not show any significant 

decrease in margin for the NEXT thruster. At best we can consider the models to be conservative with 

their predictions of lifetime prior to backstreaming, but the degree of uncertainty requires additional work 

to assure backstreaming does not occur prematurely at other throttling points.

The life-limiter of the discharge cathode is expected to be the impregnated insert. Based on planar 

impregnated cathodes, the insert is predicted to last over 70 khr at high power, which is about 1450 kg Xe

throughput and is 3.2 times the qualification requirement.

The other component that wears on the discharge cathode assembly is the keeper. The NEXT PM 

thruster will be using a graphite keeper, compared to the previous molybdenum keepers, and the wear rate 

is dramatically reduced. It is predicted that it will be well over 2000 kg of throughput before the cathode 

heater and orifice plate are significantly exposed. This prediction is corroborated by the minimal wear on 

the downstream face of the HiPEP graphite keeper after its 2000 hr test and the NEXT LDT results that 

show no measurable wear to date of the keeper orifice.

The neutralizer is anticipated to last 77 khr at high power, which is 1600 kg of Xe throughput, which 

is 3.6 times the 450 kg throughput qualification requirement. The neutralizer impregnated insert is 
NASA/TM—2010-216915 25

expected to be the life-limiter for the neutralizer cathode. It is expected that the neutralizer keeper outer 

radius wear rates on PM thrusters will be significantly reduced from the EM thruster that was tested for 

2000 hr. It is unlikely, but based on a conservative estimate, a hole could form on the keeper side from 

beam ions after 850 kg of propellant throughput, but this still wouldn’t constitute a thruster failure. One 

uncertainty for the neutralizer is orifice clogging. However, this was only documented to have occurred

during the NSTAR ELT test. This shouldn’t be an issue for the NEXT thruster since the orifice is almost 

twice the diameter of the NSTAR orifice and operates at higher emission currents. 

Additional components such as the high voltage propellant isolators and cathode heaters are either 

being tested or will be tested to verify lifetime requirements. Electrical shorting of the optics is controlled 

through retention of sputtered material in the discharge chamber, discharge chamber design minimizing 

the size of any material sputter from its walls, and the use of thermal cycling or discharge power supply to 

clear a known short. The magnets will exhibit a lifetime in excess of the NSTAR thruster magnets. After 

30,000+ hr of operation the magnets on the NSTAR ELT showed no degradation. The magnets will 

provide an adequate magnetic field beyond the thruster’s first anticipated failure mode at 750 kg.

Conclusion

Lifetime and propellant throughput capability for the NEXT ion thruster and its components based on 

several models and testing have been presented. The models include evaluation of ion optics wear, 

discharge and neutralizer cathode wear and insert depletion of low work function material, sputtering that 

could result in an unclearable high voltage short, evaluation of magnet temperatures, and voltage 

degradation of propellant isolators. The current assessment predicts that the earliest failure above a 

specific impulse of 2000 sec would be a structural failure of the accelerator grid after 750 kg of xenon 

throughput, 300 kg beyond or 1.6 to 1.7 times the 450 kg qualification requirement. Below a specific 

impulse of 2000 sec, the structural failure of the accelerator grid is predicted with less than 750 kg of 

propellant throughput, with the two lowest power throttle points below the 450 kg propellant throughput 

requirement. However, based on previous mission analysis, accelerator grid downstream erosion on the 

missions that used these low power throttle points would exhibit a maximum erosion of 51 percent

through the accelerator grid thickness. This demonstrates that while the low throttle points individually 

have propellant throughput capability below 450 kg, the throughput capability is at least twice the mission 

required. Other failure modes presented are predicted to be in excess of 750 kg thruster xenon throughput.

Modeling has not correctly predicted how electron backstreaming will change with time, but fortunately, 

the testing at the peak throttle point has shown little change. Additional testing and modeling is necessary 

to improve the fidelity of this assessment. However, even without this additional testing, there is ample 

margin to assure the NEXT thruster will meet its qualification requirement of 450 kg.

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35. Britton, M., Soulas, G., Kamhawi, H., Snyder, A., “Destructive Analysis of the NEXT 2000-Hour 

Wear Test Hollow Cathode Assemblies,” National Aeronautics and Space Administration Technical 

Memo, NASA/TM—2005-213387, Jul. 2005.

36. Kamhawi, H., Soulas, G.C., Patterson, M.J., and Frandina, M.M., “NEXT Ion Engine 2000 hour 

Wear Test Plume and Erosion Results,” 40th Joint Propulsion Conference and Exhibit, AIAA–2004–

3792, Fort Lauderdale, FL, Jul. 11–14, 2004.

37. Van Noord, J.L., Kamhawi, H., McEwen, H.K., “Characterization of a High Current Long Life 

Hollow Cathode,” The 29th International Electric Propulsion Conference, IEPC–2005–321, Princeton, 

NJ, Oct. 31–Nov. 4, 2005.

38. Kuharski, R.A., Mandell, M.J., Gardner, B.M., Katz, I., and Vaughan, D., “Ion Engine Neutralizer 

Erosion in Lab and Space,” 41st Joint Propulsion Conference and Exhibit, AIAA–2005–3880, Tucson, 

AZ, Jul. 10–13, 2005.

39. Kovaleski, S.D., Patterson, M.J., Soulas, G.C., and Sarver-Verhey, T.R., “A Review of Testing 

Hollow Cathodes for the International Space Station Plasma Contactor,” 27th International Electric 

Propulsion Conference, IEPC–01–271, Pasadena, CA, Oct. 14–19, 2001.

40. Patterson, M.J., Verhey, T.R., Soulas, G., and Zakany, J., “Space Station Cathode Design, 

Performance, and Operating Specifications,” 25th International Electric Propulsion Conference, 

IEPC–97–170, Aug. 24–28, 1997.

41. Piñero, L.R. and Sovey, J.S., “NEXT Thruster Component Verification Testing,” 43rd Joint 

Propulsion Conference and Exhibit, AIAA–2007–5276, Cincinnati, OH, Jul. 8–11, 2007.

42. Soulas, G.C., “Improving the Total Impulse Capability of the NSTAR Ion Thruster with ThickAccelerator-Grid Ion Optics,” 27th International Electric Propulsion Conference, IEPC–01–081, 

Pasadena, CA, Oct. 14–19, 2001.
NASA/TM—2010-216915 28

43. Soulas, G.C., Haag, T.W., and Patterson, M.J., “Performance Evaluation of 40 cm Ion Optics for the 

NEXT Ion Engine,” 38th Joint Propulsion Conference and Exhibit, AIAA–2002–3834, Indianapolis, 

IN, Jul. 7–10, 2002.

44. Anderson, J.R., Katz, I., and Goebel, D., “Numerical Simulation of Two-Grid Ion Optics Using a 

3D Code,” 40th Joint Propulsion Conference and Exhibit, AIAA–2004–3782, Fort Lauderdale, FL, 

Jul. 11–14, 2004.

45. Malone, S.P. and Soulas, G.C., “Computational Ion Optics Design Evaluation,” 40th Joint Propulsion 

Conference and Exhibit, AIAA–2004–3784, Fort Lauderdale, FL, Jul. 11–14, 2004.

46. Diaz, E.M. and Soulas, G.C., “Grid Gap Measurement for NSTAR Ion Thruster,” The 29th

International Electric Propulsion Conference, IEPC–2005–244, Princeton, NJ, Oct. 31–Nov. 4, 2005.

47. Monheiser, J.M. and Wilbur, P.J., “Effect of Design and Operating Conditions on Accelerator-Grid 

Impingement Current,” 23rd International Electric Propulsion Conference, IEPC–93–174, Seattle, 

WA, Sep. 13–16, 1993.

48. Pullins, S., Dressler, R.A., Chiu, Y.–H., and Levandier, D.J., “Ion Dynamics in Hall Effect and Ion 

Thrusters: Xe+ + Xe Symmetric Charge Transfer,” 38th Aerospace Sciences Meeting & Exhibit, 

AIAA–2000–0603, Reno, NV, Jan. 10–13, 2000.

49. Duchemin, O.B., “An Investigation of Ion Engine Erosion by Low Energy Sputtering,” Ph.D. 

Dissertation, California Institute of Technology, Pasadena, CA, 2001.

50. Soulas, G.C. and Patterson, M. J, “NEXT Ion Thruster Performance Dispersion Analysis,” 43rd Joint 

Propulsion Conference and Exhibit, AIAA–2007–5213, Cincinnati, OH, Jul. 8–11, 2007.

51. Anderson, J.R., Sengupta, A., and Brophy, J.R., “Post-Test Analysis of the Deep Space One Spare 

Flight Thruster Ion Optics,” 40th Joint Propulsion Conference and Exhibit, AIAA–2004–3610, Fort 

Lauderdale, FL, Jul. 11–14, 2004.

52. Hickman, T.A., et al., “Overview of Diagnostics for the NEXT Long Duration Test,” 41st Joint 

Propulsion Conference and Exhibit, AIAA–2005–4064, Tucson, AZ, Jul. 10–13, 2005.

53. Malone, S.P., “Investigation of NEXT Ion Optics Erosion Processes Using Computational 

Modeling,” 52nd Joint Army-Navy-NASA-Air Force (JANNAF) Propulsion Meeting (JPM), 

Monterey, CA, Dec. 2005.

54. Sovey, J.S., Dever, J.A., and Power, J.L., “Retention of Sputtered Molybdenum on Ion Engine 

Discharge Chamber Surfaces,” 27th International Electric Propulsion Conference, IEPC–01–086, 

Pasadena, CA, Oct. 14–19, 2001.

55. Power, J.L. and Hiznay, D.J., “Solutions for Discharge Chamber Sputtering and Anode Deposit 

Spalling in Small Mercury Ion Thruster,” NASA TM X–71675, Mar. 1975.

56. Polk, J.E., et al., “Performance of the NSTAR Ion Propulsion System on the Deep Space One 

Mission,” 39th AIAA Aerospace Sciences Meeting, AIAA–2001–0965, Reno, NV, Jan. 8–11, 2001.

57. Goodfellow, K.D., Ganapathi, G.B., and Stocky, J.F., “An Experimental and Theoretical Analysis of 

the Grid Clearing Capability of the NSTAR Ion Propulsion System,” 35th Joint Propulsion 

Conference and Exhibit, AIAA–1999–2859, Los Angeles, CA, Jun. 20–24, 1999.

58. Anderson, J.R., Vaughn, D., and Fitzgerald, D., “Experimental and Theoretical Analysis for 

Designing a Grid Clearing System for the NEXT Ion Propulsion System,” 41st Joint Propulsion 

Conference and Exhibit, AIAA–2005–3866, Tucson, AZ, Jul. 10–13, 2005.

59. Van Noord, J., “NEXT Ion Thruster Thermal Model,” 43rd Joint Propulsion Conference and Exhibit, 

AIAA–2007–5218, Cincinnati, OH, Jul. 8–11, 2007.

60. Christensen, J.A., et al., “Design and Fabrication of a Flight Model 2.3 kW Ion Thruster for the Deep 

Space 1 Mission,” 34th Joint Propulsion Conference and Exhibit, AIAA–1998–3327, Jul. 1998.

61. Sovey, J.S., Zakany, J.S. and Manzella, D.H., “NSTAR High Voltage Propellant Isolator Test,” 

NASA/TM—2004-212921, Feb. 2004.

62. Hoskins, W.A., et al., “Development of a Prototype Model ion Thruster for the NEXT System,” 40th 

Joint Propulsion Conference and Exhibit, AIAA–2004–4111, Fort Lauderdale, FL, Jul. 11–14, 2004.

63. Hoskins, W.A., et al., “Overview of the NEXT Ion Propulsion System Program at Aerojet,” 41st Joint 

Propulsion Conference and Exhibit, AIAA–2005–3885, Tucson, AZ, Jul. 10–13, 2005.
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4. TITLE AND SUBTITLE

Lifetime Assessment of the NEXT Ion Thruster 

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6. AUTHOR(S)

Van Noord, Jonathan, L. 

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13. SUPPLEMENTARY NOTES

14. ABSTRACT

Ion thrusters are low thrust, high specific impulse devices with required operational lifetimes on the order of 10,000 to 100,000 hr. The NEXT ion thruster is the latest 

generation of ion thrusters under development. The NEXT ion thruster currently has a qualification level propellant throughput requirement of 450 kg of xenon, which 

corresponds to roughly 22,000 hr of operation at the highest throttling point. Currently, a NEXT engineering model ion thruster with prototype model ion optics is undergoing a 

long duration test to determine wear characteristics and establish propellant throughput capability. The NEXT thruster includes many improvements over previous generations 

of ion thrusters, but two of its component improvements have a larger effect on thruster lifetime. These include the ion optics with tighter tolerances, a masked region and 

better gap control, and the discharge cathode keeper material change to graphite. Data from the NEXT 2000 hr wear test, the NEXT long duration test, and further analysis is 

used to determine the expected lifetime of the NEXT ion thruster. This paper will review the predictions for all of the anticipated failure mechanisms. The mechanisms will 

include wear of the ion optics and cathode’s orifice plate and keeper from the plasma, depletion of low work function material in each cathode’s insert, and spalling of material 

in the discharge chamber leading to arcing. Based on the analysis of the NEXT ion thruster, the first failure mode for operation above a specific impulse of 2000 sec is expected 

to be the structural failure of the ion optics at 750 kg of propellant throughput, 1.7 times the qualification requirement. An assessment based on mission analyses for operation 

below a specific impulse of 2000 sec indicates that the NEXT thruster is capable of double the propellant throughput required by these missions. 

15. SUBJECT TERMS

Ion thruster; Ion optics; Hollow cathodes 

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