Michael M. Frandina and Lynn A. Arrington

QSS Group, Inc., Cleveland, Ohio

George C. Soulas, Tyler A. Hickman, and Michael J. Patterson

Glenn Research Center, Cleveland, Ohio

Status of the NEXT Ion Thruster Long

Duration Test

NASA/TM—2005-213869

September 2005

AIAA–2005–4065
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Michael M. Frandina and Lynn A. Arrington

QSS Group, Inc., Cleveland, Ohio

George C. Soulas, Tyler A. Hickman, and Michael J. Patterson

Glenn Research Center, Cleveland, Ohio

Status of the NEXT Ion Thruster Long

Duration Test

NASA/TM—2005-213869

September 2005

National Aeronautics and

Space Administration

Glenn Research Center

Prepared for the

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cosponsored by the AIAA, ASME, SAE, and ASEE

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AIAA–2005–4065
Acknowledgments

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of constructing a new vacuum facility along with fabricating the necessary support equipment

required for the NEXT Long Duration Test (LDT).

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NASA/TM—2005-213869 1

Status of the NEXT Ion Thruster Long Duration Test 

Michael M. Frandina and Lynn A. Arrington 

QSS Group, Inc. 

Cleveland, Ohio 44135 

George C. Soulas, Tyler A. Hickman, and Michael J. Patterson 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Abstract 

The status of NASA’s Evolutionary Xenon Thruster (NEXT) Long Duration Test (LDT) is presented. The test 

will be conducted with a 36 cm diameter engineering model ion thruster, designated EM3, to validate and qualify 

the NEXT thruster propellant throughput capability of 450 kg xenon. The ion thruster will be operated at various 

input powers from the NEXT throttle table. Pretest performance assessments demonstrated that EM3 satisfies all 

thruster performance requirements. As of June 26, 2005, the ion thruster has accumulated 493 hours of operation 

and processed 10.2 kg of xenon at a thruster input power of 6.9 kW. Overall ion thruster performance, which 

includes thrust, thruster input power, specific impulse, and thrust efficiency, has been steady to date with very little 

variation in performance parameters. 

I. Introduction 

The success of the NASA Solar Electric Propulsion Technology Applications Readiness (NSTAR) program’s 

ion propulsion system on the Deep-Space 1 spacecraft has secured the future for ion propulsion technology for other 

NASA missions (refs. 1 and 2). While the 2.3 kW NSTAR ion thruster input power and service life capabilities are 

appropriate for Discovery Class as well as other, smaller NASA missions, the application of NSTAR hardware to 

more demanding missions such as outer planet explorers and sample return missions is limited due its lack of power 

and total impulse capability. As a result, NASA’s Office of Space Science awarded a development project to a 

NASA Glenn Research Center (GRC)-led team to develop the next generation ion propulsion system (refs. 3 and 4). 

The propulsion system, called NASA’s Evolutionary Xenon Thruster (NEXT), consists of a 36 cm beam extraction 

diameter ion thruster, a lightweight, modular power processing unit with an efficiency and a specific mass equal-to 

or better-than the NSTAR power processor, and a xenon feed system that will significantly reduce mass and volume 

relative to the NSTAR feed system. Ion thruster performance requirements include a specific impulse of at least 

4050 s at full power and a thruster efficiency of greater than 0.63 at full power. Regarding service life, the ion 

thruster must provide a 185 kg baseline requirement and a 300 kg nominal mission requirement xenon throughput 

capability. Therefore, based on the nominal mission requirement, the ion thruster must provide a 450 kg 

qualification throughput (ref. 5). The first wear test of a NEXT ion thruster was for a 2038 hour duration at full 

power and processed 43 kg of xenon (ref. 6). The service life capability of the NEXT ion thruster is being assessed 

by thruster wear tests and life-modeling of critical thruster components, such as the ion optics and cathodes. 

The maximum thruster input power level was chosen for the first 300 kg of propellant throughput because lifemodeling predicted that this power level will cause the most severe thruster erosion. The objectives of the LDT 

include: characterizing thruster operation and performance over the duration of the test; identifying thruster lifelimiting phenomena; and measuring thruster component wear rates and comparing them with that predicted from life 

models. In addition, after 300 kg propellant throughput has been achieved the thruster will be throttled into a design 

reference mission profile until qualification level propellant throughput has been demonstrated (ref. 5). This paper 

presents the status of the NEXT LDT. A description of the test article is discussed followed by a description of the 

test support hardware, which includes the power console, gas feed system, vacuum facility, and diagnostics. The test 

operating condition is then described. Finally, wear test results to date are presented and discussed. 

II. Test Article 

The wear test is being conducted with an engineering model ion thruster, designated EM3, which is shown in 

figure 1. The thruster utilizes a 36 cm beam extraction diameter, which yields approximately 1.6 times the area of 

the NSTAR ion thruster. The technical approach here is a continuation of the “derating” philosophy used for the 
NASA/TM—2005-213869 2

NSTAR ion thruster. A beam extraction area 1.6 times that of the NSTAR ion thruster beam area allows operation at 

significantly higher thruster input power while maintaining low voltages and ion current densities. Thus, potential 

complications associated with high voltage electrode operations are avoided, and thruster longevity can be 

maintained. The discharge chamber design of EM3 utilizes a hollow cathode electron emitter and a semi-conic 

chamber with a ring cusp magnetic circuit for electron containment created by high strength, rare earth magnets. A 

flake retention scheme is employed in the discharge chamber, which also acts as a magnet retainer. The material, 

preparation, and installation processes employed for the flake-retention system are identical to those implemented on 

the NSTAR thruster (ref. 7). The discharge chamber also incorporates a reverse-feed propellant injection process for 

the main plenum. Finally, EM3 utilizes new compact propellant isolators with a higher voltage isolation capability 

(about 1800 V) than those used by the NSTAR thruster. The ion thruster utilizes a neutralizer design that is 

mechanically similar to the Hollow Cathode Assembly of the International Space Station Plasma Contactor (ref. 8). 

Because the neutralizer cathode emission current range on the NEXT ion thruster is similar to that of the Plasma 

Contactor Hollow Cathode Assembly, the NEXT neutralizer design can leverage the large cathode database already 

available with this design for risk reduction (ref. 9). The ion optics’ are a set of prototype model hardware 

manufactured by Aerojet (ref. 10). The mounting assembly of EM3 is similar to that of the NEXT ion thruster EM1, 

which was used in the NEXT 2000 Hour Wear Test (ref. 6). The performance of the EM1 ion optics electrode 

geometry has been described in detail (ref. 11). 

III. Test Support Hardware 

A. Power Console 

A power console similar to that described in reference 12 and used in reference 6 powers the ion thruster. This 

power console utilizes commercially available power supplies that allow for ion thruster input powers of over 10 kW 

with beam power supply voltages of up to 2000 V. 

B. Gas Feed System 

A high purity gas feed system is used to provide xenon to the discharge cathode and neutralizer cathode through 

separate mass flow controllers. For the main plenum, xenon is provided either through a separate mass flow 

controller or through an engineering model propellant management system kernel provided by Aerojet (ref. 13). 

C. Vacuum Facility 

The wear test is being conducted in Vacuum Facility 16 at NASA GRC, shown in figure 2. This 2.7 m diameter 

8.5 m long facility has a bell jar on the end cap, which can be isolated from the main body of the facility. In case of 

an emergency the thruster could be withdrawn into the bell jar. Vacuum Facility 16 is evacuated with 10 cryogenic 

pumps, with an additional cryogenic pump on the isolated bell jar for emergency use. For this test, the facility 

pressure is monitored by two ionization gages, with one located in the facility about 0.5 m radially beside the ion 

thruster and the second located on the cylindrical section of the facility wall about 0.5 m downstream of the ion 

thruster. In addition, the isolated bell jar has an ion gage that is turned off during normal operation. The facility base 

pressure has been less than 3.3*10–5 Pa (2.5*10–7 Torr). Facility background pressure during thruster operation at 

full power has been 3.3*10–4 Pa (2.5*10–6 Torr) next to the ion thruster. The facility pumping speed on xenon 

calculated with this ion gage is approximately 170,000 L/s. The facility interior surfaces downstream of the ion 

thruster are lined with a carbon material to reduce the amount of facility material that is backsputtered onto the ion 

thruster. The backsputter rate is determined with a quartz crystal microbalance located next to the ion thruster. The 

backsputter rate during wear testing has typically been about 3 microns/kh. 

D. Diagnostics 

A computerized data acquisition and control system is used to monitor ion thruster and facility operations. Data 

are sampled at a frequency range of 10 to 20 Hz and is stored once per minute. Ion thruster currents and voltages are 

measured with current shunts and voltage dividers, respectively, and recorded. Facility pressures and individual 

mass flows to the ion thruster are also sampled and recorded. In addition, the thruster is periodically connected to an 

electrically floating power supply circuit used to determine the screen grid transparency to discharge chamber ions. 

The circuit electrically ties the screen grid to the discharge cathode during normal operation, but during this 

diagnostic the grid is biased negative relative to discharge cathode potential to repel electrons and measure the 

collected ion current. 

Ion beam diagnostics and erosion diagnostics are also included in this test and are described in a companion 

paper (ref. 14). 
NASA/TM—2005-213869 3

IV. Operating Condition and Performance Tests

The NEXT ion thruster is designed for solar electric propulsion applications. Ion thruster input power is, 

therefore, designed to be throttled from 0.5 to 6.9 kW. For the LDT, it is necessary to demonstrate the thruster’s 

propellant throughput capability at an operating point that causes the most severe and stressful thruster operation in 

addition to demonstrating performance at several throttle table operating points. The highest thruster input power, 

which corresponded to a 1800 V beam power supply voltage and a 3.52 A beam current, was therefore selected as 

the wear test operating condition for the first 300 kg propellant throughput. The NEXT LDT thruster full power 

operating point is listed in table 1. Prior to and throughout the wear test, performance tests are conducted on the 

thruster and thruster components. Thruster performance tests include measuring thruster operating parameters and 

determining thruster performance at several of the power levels in the throttle table. Component performance 

assessments are periodically made on the discharge chamber, ion optics, and neutralizer cathode. Ion optics 

performance includes electron backstreaming, perveance, and screen grid ion transparency measurements (ref. 15). 

Discharge chamber performance is assessed by measuring discharge losses as a function of discharge propellant 

utilization efficiency at fixed discharge voltages. Finally, neutralizer performance, which included keeper voltage 

measurements as a function of neutralizer flow, is assessed. 

V. Test Results and Discussions

As of June 26, 2005, NEXT thruster EM3 has accumulated 493 hours of operation at a thruster input power of 

6.9 kW. There have been a total of 4 unintended test interruptions; however, none were caused by abnormal thruster 

operation. 

A. Pretest Performance Results 

Table 2 lists some of the pretest ion thruster performance results along with the beginning of life throttle table 

values at various operating conditions. For thrust calculations in table 2, the beam divergence thrust correction factor 

and the total doubly-to-singly-charged ion current ratio were assumed to be in the 0.962 to 0.974 and 0.028 to 0.060 

ranges, respectively, based on the methodology developed for NSTAR thrusters (ref. 16). Ingested mass flow due to 

the facility background gas pressure was included in the total mass flow rate to the thruster for determining thrust 

efficiency and specific impulse (ref. 17). The demonstrated throttling range was 0.5 to 6.9 kW, with a resulting 

thrust of about 26 mN at low power to 237 mN at the wear test full power point. The corresponding specific impulse 

range was 1360 to 4160 s. Thrust efficiencies were between about 0.319 at low power to 0.706 at the wear test full 

power point. 

Along with thruster performance data, pretest discharge chamber and neutralizer performance were also 

determined. Figure 3 shows discharge losses as a function of discharge chamber propellant utilization efficiency for 

all throttle table and wear test beam currents. At each beam current, the ratio of main to discharge cathode flow rate 

was adjusted to maintain the nominal discharge voltage. Ingested mass flow due to the facility background gas 

pressure was included in the calculation of discharge propellant utilization efficiency. All data, except for the lowest 

power point, were obtained for a beam power supply voltage of 1179 V. This voltage was chosen both because it 

was the lowest voltage that was common to all beam currents in table 1, except the lowest power point, and it also 

provided the highest discharge losses. At each of these beam currents, discharge losses were recorded between 

utilization efficiencies of about 0.85 to 0.95 when possible. As shown in figure 2, discharge losses at beam currents 

between 2.70 A and 3.52 A were within 10 W/A throughout all propellant utilization efficiencies examined. 

Discharges losses increased more significantly for the 1.00, 1.20 and 2.00 A beam currents. 

Figure 4 shows neutralizer keeper voltage as a function of flow rate for several beam currents for both EM3 and 

EM1 (ref. 18). Data were measured at a beam power supply voltage of 1800 V. The lowest flow value for each 

curve was either in plume mode, which is defined here as a peak-to-peak neutralizer keeper voltage oscillation of 

greater than 5 V, or was the last point taken due to a power supply failure. For all neutralizer flow rates and beam 

currents (Jb), EM3 has a higher neutralizer voltage of up to about 1 V. The figure indicates that for the 4.01 sccm 

neutralizer flow for the full power LDT operating condition there is at least 2.0 sccm flow margin between the 

nominal operating flow and plume mode operation. 

B. Wear Test Results 

The LDT began when full power operation was first initiated. During the first few hours, performance tests were 

conducted at the full power point. For all wear test results presented, data recorded during performance testing have 

been removed to demonstrate the thruster’s behavior as function of time, not as a function of varying input values. 
NASA/TM—2005-213869 4

1. Thruster Performance 

Thruster performance parameters of value to mission planners include thrust, input power, efficiency, and 

specific impulse. Thrust and input power as a function of time are plotted in figures 5 and 6, respectively. Thrust 

was determined as described in the prior section. Thrust has remained steady at 237 mN ± 2 mN. Input power has 

remained steady at 6.85 kW ± 0.05 kW. 

Specific impulse and thrust efficiency as a function of time are plotted in figures 7 and 8, respectively. All 

thruster mass flow rates, which are used to calculate thrust efficiency and specific impulse, have been within 

± 1 percent of their nominal set points since hour 94 of the LDT. Before hour 94, the neutralizer cathode flow rate 

was set at a value that was higher than the throttle table value. Since then, specific impulse and thrust efficiency 

have been 4180 s ± 30 s and 0.710 ± 0.005, respectively. Variations in performance have been less than ± 1 percent 

of the nominal values. 

2. Discharge Chamber 

Discharge current and voltage as a function of time are shown in figures 9 and 10, respectively. The discharge 

current is adjusted as needed to maintain a constant beam current of 3.52 A ± 0.01 A. As the figure shows, changes 

with time have generally been less than 3 percent of the nominal run values. The only noted trend, while small, has 

been with the discharge current, which rose from a nominal start value of about 18.5 A to about 18.9 A by hour 350. 

Discharge voltages to date have been 23.2 V +0.4/–0.3 V. Discharge losses, shown in figure 11, increased from 

123 W/A to about 124 W/A by hour 350. This change is within the measurement uncertainty. The relatively constant 

discharge losses over 350 hours is contrasted with that exhibited by the NSTAR thruster. In three separate wear 

tests, NSTAR thruster discharge losses increased by 10 to 15 W/A within the first 500 hours of wear testing (refs. 19 

to 21). The initial change in the NEXT thruster discharge losses was very small and this contributes to a more 

constant thruster input power, and, therefore, a more constant thrust efficiency at beginning of life. 

3. Neutralizer 

Neutralizer keeper current as a function of time is plotted in figure 12. Neutralizer keeper currents to date have 

been 3.00 A ± 0.02 A. Neutralizer keeper voltage as a function of time is plotted on figure 12. Since the neutralizer 

flow was adjusted after hour 94, the neutralizer keeper voltages to date have been 11.1 V +0.1/–0.3 V. The 

neutralizer keeper voltage exhibits voltage “spikes”, as shown in figure 13. These voltage “spikes” correspond to 

thruster restarts, which is similar to NSTAR neutralizer behavior (refs. 16, 17, and 19). Coupling voltage, which is a 

measure of neutralizer cathode potential relative to vacuum facility ground, is plotted in figure 14 as a function of 

time. The coupling voltage has been steady since hour 94 at –10.2 V +0.3/–0.1 V. The low coupling voltage 

magnitude is due to the high keeper current and the considerable neutralizer flow margin selected for operation. 

These high values were selected to ensure neutralizer operation in spot mode throughout ion thruster service life 

while imposing only modest sacrifices in thruster performance. Both the neutralizer keeper and coupling voltage 

variations indicate no degradation of neutralizer performance to date. 

4. Ion Optics 

Accelerator grid current as a function of time is plotted in figure 15. The accelerator current decreased from a 

start value of about 16 mA to about 14.2 mA by hour 94. The cause of the decrease is threefold. First, from test start 

to hour 49 the vacuum facility had only 9 of the 10 cryopumps operating which caused a higher background pressure 

that in turn caused a higher accelerator current. Second, the thruster experienced a burn in period that is typical to 

NEXT thrusters. Finally, at hour 94 the neutralizer flow was reduced which caused a lower background pressure that 

in turn caused a lower accelerator current. Since hour 94 the accelerator currents have been steady at 14.2 mA 

+0.5/–0.2 mA. The “spikes” in the accelerator current correlate with thruster recycles. Immediately following a 

thruster recycle, the accelerator current returns to an increased value. The current then decreases gradually for up to 

10 minutes to its nominal value. This behavior may be the result of grid cooling during a recycle. During a thruster 

recycle, the discharge power is momentarily reduced so that when high voltage is re-applied, another recycle is not 

inadvertently triggered. It is speculated that this momentary reduction in discharge power allows the ion optics to 

cool enough to change the hot grid gap, thus changing the accelerator current by a small amount. Regardless of the 

cause, these accelerator current “spikes” are not considered harmful to thruster operation. Beginning-of-life 

accelerator currents for the NEXT thruster were about 13 percent higher than nominal accelerator currents and 

reached nominal values within about 94 hours. NSTAR thruster accelerator currents generally started higher-thannominal and required up to 1500 hours to decrease to nominal values (refs. 20 and 21). Accelerator voltage as a 

function of time is plotted in figure 16. The accelerator voltages have been steady at –210 V ± 2 V. 

Correcting thruster recycles by removing those occurring during performance testing yields 283 total thruster 

recycles as of hour 493. The recycle rate as a function of time is plotted in figure 17. The recycle rate has always 

been less than 1.5 recycles per hour. In comparison, the NSTAR 8200 hour wear test operated with an average 

recycle rate of 1 to 2 per hour (ref. 20). It was expected that the NEXT thruster should have a slightly higher average 
NASA/TM—2005-213869 5

recycle rate than the NSTAR thruster because the NEXT thruster’s beam extraction area and inter-grid electric field 

are 1.6X and 1.5X that of the NSTAR thruster, respectively. Impingement-limited total voltages, electron 

backstreaming limits, and screen grid ion transparencies throughout the wear test are plotted in figures 18, 19, and 

20, respectively. Impingement-limited total voltages were determined from plots of accelerator current as a function 

of total ion extraction voltage where the slope was –0.02 mA/V (ref. 15). Electron backstreaming limits were 

determined by lowering the magnitude of the accelerator grid voltage until the indicated beam current increased by 

1 mA due to backstreaming electrons. Screen grid ion transparencies were calculated in the same manner as 

previous NEXT thruster testing (ref. 11). To date, impingement limited total voltages have been steady at 830 to 

880 V, electron backstreaming limits have been steady at –176 to –175 V, and screen grid ion transparencies have 

been steady at 0.89 to 0.90. All three of these performance parameters indicate no significant ion optics’ 

performance degradation. 

VI. Conclusion 

The LDT is being conducted with a 36 cm, engineering model NEXT ion thruster EM3. As of June 26, 2005, 

NEXT thruster EM3 has accumulated about 493 hours of operation at a thruster input power of 6.9 kW. Pretest ion 

thruster performance was determined over a throttling range of 0.5 to 6.9 kW, with resulting thrusts of about 26 mN 

at low power to 237 mN at the full power point. The corresponding specific impulse range was 1360 to 4160 s. 

Thrust efficiencies were between about 0.319 at low power to 0.706 at full power. Pretest results demonstrated that 

EM3 satisfies all thruster performance requirements. Discharge chamber, neutralizer, and ion optics operation 

during the wear test was evaluated. Discharge losses increased by less than 1 percent during the first 493 hours. 

Neutralizer keeper and coupling voltages have been steady, with no indication of performance degradation. 

Impingement limited total voltages, electron backstreaming limits, and screen grid ion transparencies have also been 

steady, with no indication of performance degradation. Overall ion thruster performance, which includes thrust, 

thruster input power, specific impulse, and thrust efficiency, has been steady to date. Variations in performance have 

been less than ± 1 percent of nominal values. 

References 

1. Rawlin, V.K., et al., “An Ion Propulsion System for NASA’s Deep Space Missions,” AIAA Paper 99–4612, 

September 1999. 

2. Polk, J.E., et al., “In-Flight Performance of the NSTAR Ion Propulsion System on the Deep Space One 

Mission,” IEEE Aerospace Conference Proceedings 8.0304, March 2000. 

3. Patterson, M.J., et al., “NEXT: NASA's Evolutionary Xenon Thruster Development Status,” AIAA Paper 2003–

4862, July 2003. 

4. Patterson, M.J., et al., “NEXT: NASA's Evolutionary Xenon Thruster,” AIAA Paper 2002-3832, July 2002. 

5. Oh, D., et al, “Deep Space Mission Applications for NEXT: NASA’s Evolutionary Xenon Thruster,” AIAA 

Paper 2004–3806, July 2004. 

6. Soulas, G.C., et al., “NEXT Ion Engine 2000 Hour Wear Test Results,” AIAA Paper 2004-3791, July 2004. 

7. Christensen, J.A., et al., “Design and Fabrication of a Flight Model 2.3 kW Ion Thruster for the Deep Space 1 

Mission,” AIAA Paper 98–3327, July 1998. 

8. Patterson, M.J., et al., “Space Station Cathode Design, Performance, and Operating Specifications,” IEPC Paper 

97–170, August 1997. 

9. Kovaleski, S.K., et al., “A Review of Testing of Hollow Cathodes for the International Space Station Plasma 

Contactor,” IEPC Paper 01–271, October 2001. 

10. Hoskins, W., “Design and Fabrication of the NEXT Prototype Model Ion Thruster,” AIAA Paper 2005–3885, 

July 2005. 

11. Soulas, G.C., et al., “Performance Evaluation of 40 cm Ion Optics for the NEXT Ion Thruster,” AIAA Paper 

2002–3834, July 2002. 

12. Pinero, L.R., et al., “Power Console Development for NASA’s Electric Propulsion Outreach Program,” IEPC 

Paper 93–250, September 1993. 

13. Aadland, R.S., et al., “Development Status of the NEXT Propellant Management System,” AIAA Paper 2004–

3974, July 2004. 

14. Hickman, T.A., et al., “Overview of Diagnostics for the NEXT Long Duration Test,” AIAA Paper 2005–4064, 

July 2005. 

15. Soulas, G.C., et al., “Performance Evaluation of the NEXT Ion Engine,” AIAA Paper 2003–5278, July 2003. 

16. Patterson, M.J., et al., “Performance of the NASA 30 cm Ion Thruster,” IEPC Paper 93–108, September 1993. 
NASA/TM—2005-213869 6

17. Sovey, J.S., “Improved Ion Containment Using a Ring-Cusp Ion Thruster,” Journal of Spacecraft and Rockets, 

vol. 21, no. 5, September-October 1984, p. 489. 

18. Soulas, G.C., et al., “Status of the NEXT Ion Engine Wear Test,” AIAA Paper 2003-4863, July 2003. 

19. Polk, J.E., et al., “A 1000-Hour Wear Test of the NASA NSTAR Ion Thruster,” AIAA Paper 96–2717, July 

1996. 

20. Polk, et al., “An Overview of the Results from an 8200 Hour Wear Test of the NSTAR Ion Thruster,” AIAA 

Paper 99–2446, June 1999. 

21. Anderson, J.R., et al., “Results of an On-going Long Duration Ground Test of the DS1 Flight Spare Ion 

Thruster,” AIAA Paper 99–2857, June 1999. 

TABLE 1.—NEXT LDT OPERATING POINT 

TABLE 2.—NEXT THRUSTER EM3 PRETEST PERFORMANCE RESULTS AT 

VARIOUS BEGINNING OF LIFE THROTTLE TABLE OPERATING CONDITIONS 

Figure 1.—NEXT engineering model thruster EM3. 

Engine Input 

Powera

, kW

Beam Current, 

A

Beam Power 

Supply 

Voltage, V

Accelerator 

Voltage, V

Main Plenum 

Flow Rateb

, 

sccm

Discharge 

Cathode Flow 

Rateb

, sccm

Neutralizer 

Cathode Flow 

Ratec, sccm

Neutralizer 

Keeper 

Current, A

6.83 3.52 1800 -210 49.64 4.87 4.01 3.00

a

Nominal Values

b

Main-to-discharge cathode flow split selected to result in a 23.5-27 V discharge voltage

Beam Current, 

A

Beam Power 

Supply 

Voltage, V

Thruster Input 

Power, kW

Discharge 

Losses, W/A

Thrust 

Efficiency Thrust, mN Specific Impulse, s

Performance test result 1.00 275 0.54 209 0.328 26 1400

Throttle table value 1.00 275 0.53 216 0.331 26 1400

Performance test result 1.20 1801 2.44 187 0.641 80 3980

Throttle table value 1.20 1800 2.43 188 0.647 80 4000

Performance test result 2.01 1800 3.97 153 0.710 134 4290

Throttle table value 2.00 1800 4.00 174 0.707 134 4310

Performance test result 3.54 1801 6.84 123 0.706 237 4160

Throttle table value 3.52 1800 6.83 124 0.710 236 4190
NASA/TM—2005-213869 7

Figure 2.—NASA Glenn Research Center Vacuum Facility 16. 

100

120

140

160

180

200

220

240

260

280

300

0.84 0.86 0.88 0.90 0.92 0.94 0.96

Discharge Propellant Utilization Efficiency (Including Mass Ingestion)

1.00 A, 26.4 V, 275 V

1.20 A, 25.0 V, 1179 V

2.00 A, 23.5 V, 1179 V

2.70 A, 23.9 V, 1179 V

3.52 A, 23.7 V, 1179 V

Figure 3.—Discharge losses as a function of discharge propellant utilization efficiency for all beam currents. 
NASA/TM—2005-213869 8

10

11

12

13

14

15

16

17

18

19

1.50 2.00 2.50 3.00 3.50 4.00 4.50 5.00 5.50 6.00

Neutralizer Flow Rate, sccm

EM3, Jb = 1.20 A

EM1, Jb = 1.20 A

EM3, Jb = 2.00 A

EM1, Jb = 2.00 A

EM3, Jb = 3.52 A

EM1, Jb = 3.52 A

Figure 4.—Neutralizer keeper voltage as a function of flow rate for EM1 and 

EM3 with all beam currents at a beam power supply voltage of 1800 V. 

200

210

220

230

240

250

260

270

280

0 100 200 300 400 500 600

Total Elapsed Time, h

Figure 5.—Calculated thrust as a function of time. 

6.00

6.20

6.40

6.60

6.80

7.00

7.20

7.40

7.60

7.80

8.00

0 100 200 300 400 500 600

Total Elapsed Time, h

Figure 6.—Thruster input power as a function of time. 
NASA/TM—2005-213869 9

3500

3600

3700

3800

3900

4000

4100

4200

4300

4400

4500

0 100 200 300 400 500 600

Total Elapsed Time, h

Figure 7.—Calculated specific impulse as a function of time. 

0.50

0.55

0.60

0.65

0.70

0.75

0.80

0 100 200 300 400 500 600

Total Elapsed Time, h

Figure 8.—Calculated thrust efficiency as a function of elapsed time. 

16

17

18

19

20

21

22

0 100 200 300 400 500 600

Total Elapsed Time, h

Figure 9.—Discharge current as a function of time. 
NASA/TM—2005-213869 10

21

22

23

24

25

26

27

0 100 200 300 400 500 600

Total Elapsed Time, h

Figure 10.—Discharge voltage as a function of time. 

100

105

110

115

120

125

130

135

140

145

150

0 100 200 300 400 500 600

Total Elapsed Time, h

Figure 11.—Discharge losses as a function of time. 

2.80

2.85

2.90

2.95

3.00

3.05

3.10

3.15

3.20

0 100 200 300 400 500 600

Total Elapsed Time, h

Figure 12.—Neutralizer keeper current as a function of time. 
NASA/TM—2005-213869 11

9.0

9.5

10.0

10.5

11.0

11.5

12.0

12.5

13.0

0 100 200 300 400 500 600

Total Elapsed Time, h

Figure 13.—Neutralizer keeper voltage as a function of time. 

-12.0

-11.5

-11.0

-10.5

-10.0

-9.5

-9.0

-8.5

-8.0

0 100 200 300 400 500 600

Total Elapsed Time, h

Figure 14.—Coupling voltage as a function of time. 

8

10

12

14

16

18

20

0 100 200 300 400 500 600

Total Elapsed Time, h

Figure 15.—Accelerator current as a function of time. 
NASA/TM—2005-213869 12

-230

-225

-220

-215

-210

-205

-200

-195

-190

0 100 200 300 400 500 600

Total Elapsed Time, h

Figure 16.—Accelerator voltage as a function of time. 

0.0

0.5

1.0

1.5

2.0

2.5

3.0

0 100 200 300 400 500 600

Total Elapsed Time, h

Corrected for recycles during performance testing

Calculated by linear regression over ~ 12 hour period

Figure 17.—Recycle rate as a function of time. 

600

650

700

750

800

850

900

950

1000

1050

1100

0 50 100 150 200 250 300 350 400

Total Elapsed Time, h

Figure 18.—Impingement limited total voltage at full power as a function of time. 
NASA/TM—2005-213869 13

-200

-195

-190

-185

-180

-175

-170

-165

-160

-155

-150

0 50 100 150 200 250 300 350 400

Total Elapsed Time, h

Vbps = 1800 V, Jb = 3.52 A

Figure 19.—Electron backstreaming limit as a function of time. 

0.80

0.82

0.84

0.86

0.88

0.90

0.92

0.94

0.96

0.98

1.00

0 50 100 150 200 250 300 350 400

Total Elapsed Time, h

Figure 20.—Screen grid ion transparency as a function of time. 
This publication is available from the NASA Center for AeroSpace Information, 301–621–0390.

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19

Status of the NEXT Ion Thruster Long Duration Test

Michael M. Frandina, Lynn A. Arrington, George C. Soulas, Tyler A. Hickman,

and Michael J. Patterson

Ion propulsion; Ion beams; Ion optics; Wear tests

Unclassified -Unlimited

Subject Category: 20

Prepared for the 41st Joint Propulsion Conference and Exhibit cosponsored by the AIAA, ASME, SAE, and ASEE,

Tucson, Arizona, July 10–13, 2005. Michael M. Frandina and Lynn A. Arrington, QSS Group, Inc., 21000 Brookpark

Road, Cleveland, Ohio 44135; and George C. Soulas, Tyler A. Hickman, and Michael J. Patterson, NASA Glenn

Research Center. Responsible person, Lynn A. Arrington, organization code RPP, 216–977–7486.

The status of NASA’s Evolutionary Xenon Thruster (NEXT) Long Duration Test (LDT) is presented. The test will be

conducted with a 36 cm diameter engineering model ion thruster, designated EM3, to validate and qualify the NEXT

thruster propellant throughput capability of 450 kg xenon. The ion thruster will be operated at various input powers from

the NEXT throttle table. Pretest performance assessments demonstrated that EM3 satisfies all thruster performance

requirements. As of June 26, 2005, the ion thruster has accumulated 493 hours of operation and processed 10.2 kg of

xenon at a thruster input power of 6.9 kW. Overall ion thruster performance, which includes thrust, thruster input power,

specific impulse, and thrust efficiency, has been steady to date with very little variation in performance parameters.