Michael J. Patterson and Scott W. Benson

Glenn Research Center, Cleveland, Ohio

NEXT Ion Propulsion System

Development Status and Capabilities

NASA/TM—2008-214988

January 2008

Paper number D10P3
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Paper number D10P3

Michael J. Patterson and Scott W. Benson

Glenn Research Center, Cleveland, Ohio

NEXT Ion Propulsion System

Development Status and Capabilities

NASA/TM—2008-214988

January 2008
Available from

NASA Center for Aerospace Information

7115 Standard Drive

Hanover, MD 21076–1320

National Technical Information Service

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Available electronically at http://gltrs.grc.nasa.gov

Level of Review: This material has been technically reviewed by technical management.
NASA/TM—2008-214988 1

NEXT Ion Propulsion System 

Development Status and Capabilities

Michael J. Patterson and Scott W. Benson 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Abstract 

NASA’s Evolutionary Xenon Thruster (NEXT) project is 

developing next generation ion propulsion technologies to 

provide future NASA science missions with enhanced mission 

performance benefit at a low total development cost. The 

objective of the NEXT project is to advance next generation 

ion propulsion technology by producing engineering model 

system components, validating these through qualificationlevel and integrated system testing, and ensuring preparedness 

for transitioning to flight system development. As NASA’s 

Evolutionary Xenon Thruster technology program completes 

advanced development activities, it is advantageous to review 

the existing technology capabilities of the system under 

development. This paper describes the NEXT ion propulsion 

system development status, characteristics and performance. A 

review of mission analyses results conducted to date using the 

NEXT system is also provided. 

I. Introduction 

The NASA Glenn Research Center (GRC) is responsible 

for the development of NEXT ion propulsion system (IPS).1

The objective of the NEXT project is to advance next 

generation ion propulsion technology to NASA Technology 

Readiness Level (TRL) 6. This two-phase project is sponsored 

by NASA Science Mission Directorate (SMD), conducted 

under the In-Space Propulsion Technology Program (ISPT). 

NEXT was implemented through a competitively-selected 

NASA Research Announcement (NRA) awarded in 2002 with 

the first phase completed in August 2003. Technology 

validation and mission analysis results in Phase 1 indicated 

that the NEXT technologies have the capabilities that provide 

the expected benefits, and further development was warranted 

(ref. 1). The second phase of the project was initiated in 

October 2003 and continues. 

The NEXT system elements (depicted in fig. 1) consist of a 

high performance, 7-kW, ion thruster; a modular, highefficiency 7-kW power processor unit (PPU); a highly flexible 

advanced xenon propellant management system (PMS) 

 

1

The NEXT team consists of NASA GRC as technology project lead, 

JPL as system integration lead, Aerojet (Redmond, WA) as PM 

thruster, PMS, and DCIU simulator developer, and L3 Comm ETI 

(Torrance, CA) as PPU developer. 

consisting of a single High Pressure Assembly (HPA) and one 

Low Pressure Assembly (LPA) per thruster; a lightweight 

engine gimbal; and key elements of a digital control interface 

unit (DCIU) including software algorithms (refs. 2 to 5). This 

design approach was selected to provide future NASA science 

missions with the greatest value in mission performance 

benefit at a low total development cost. 

The NEXT thruster and other component technologies 

represent a significant advancement in technology beyond 

state-of-the-art (SOA) NSTAR thruster systems (refs. 1 and 

6). NEXT performance exceeds single or multiple NSTAR 

thrusters over most of the thruster input power range. Higher 

efficiency and specific impulse, and lower specific mass 

reduce the wet propulsion system mass and parts count. The 

NEXT thruster xenon propellant throughput capability is more 

than twice NSTAR’s, so fewer thrusters are needed. The 

NEXT power processor and propellant feed system technologies provide mass and performance benefits versus NSTAR. 

Comparisons of NEXT and SOA NSTAR performance 

characteristics are listed in table I. The NEXT IPS development project has also placed particular emphasis on key 

aspects of IPS development with the intention of avoiding the 

difficulties experienced by the Dawn mission in transitioning 

the NSTAR-based technology to an operational ion propulsion 

system (ref. 7). These aspects are discussed in more detail 

later. 

TABLE I.—PERFORMANCE CHARACTERISTICS 

OF NEXT VERSUS NSTAR SOA 

Characteristic NEXT NSTAR SOA 

Thruster Power Range, kW 0.5-6.9 0.5-2.3 

Max. Specific Impulse, s >4100 >3100 

Max. Thrust, mN 236 92 

Max. Thruster Efficiency >70% >61% 

Max. PPU Efficiency 95% 92% 

PPU Specific Mass, kg/kW 4.8 6.0 

PMS Single-String Mass, kg 5.0 11.4 

PMS Unusable Propellant Residual 1.00% 2.40% 

NEXT technology is applicable to a wide range of NASA 

solar system exploration missions (see fig. 2), as well as earthspace commercial and other missions of national interest. 

NEXT affords larger delivered payloads and smaller launch 

vehicle size for Discovery, New Frontiers, Mars Exploration, 

and Flagship outer-planet exploration missions. 
NASA/TM—2008-214988 2

Figure 1.—NEXT Ion Propulsion Elements. 

Figure 2.—NEXT makes high-performing ion propulsion 

systems available to a wide range of NASA missions. 

II. Development Status

This section summarizes project activities conducted during 

Phase 1 and Phase 2 relative to subsystem and system 

elements, and the present and intended development status of 

the elements at Phase 2 completion. 

A. Thruster 

The NEXT thruster is a 0.54 to 6.9 kW input power, 36 cm 

beam diameter, xenon ion thruster with 2-grid ion optics. The 

beam current at full power of 6.9 kW is 3.52 A, and it has a 

maximum specific impulse of greater than 4170 s at a 

maximum thrust of greater than 236 mN, with a peak efficiency in excess of 70%. The xenon throughput requirement is 

> 300 kg, with a 450 kg qualification level. The analysis-based 

capability of the thruster is > 730 kg. 

Five engineering model (EM) thrusters have been manufactured at GRC and tested. During Phase 1 of the NEXT project 

(ending in August 2003) EM thruster testing included detailed 

performance evaluations (ref. 8), a 2000 hr wear test at full 

power (ref. 9), integration testing with a breadboard PPU and 

breadboard PMS (ref. 10) and structural characterization tests 

to characterize thruster dynamic behavior for further thruster 

development (ref. 11). During Phase 2 to date, EM thruster 

performance has been tested in a multi-thruster array (ref. 12), 

and a long-duration life test of an EM thruster has accumulated more than 9300 hr of operation at full power. 

Figure 3.—PM Thruster undergoing 

Thermal Vacuum Testing at JPL. 

Also during Phase 2 the first prototype model2

 (PM) 

thruster was manufactured by Aerojet and delivered to GRC in 

January 2006. Acceptance testing of the thruster was completed at GRC and the performance was found to be consistent 

with that demonstrated from multiple EM thrusters. The PM 

thruster was subsequently subjected to a series of validation 

tests at JPL including: thruster/gimbal functional tests, a thermal 

development test, and qualification-level environmental 

(vibration and thermal vacuum) tests (fig. 3). The thermal 

development test proceeding environmental was performed to 

develop and validate the thruster thermal model and demonstrate thruster operation and temperature margins over a large 

temperature range. The vibration test was conducted with the 

PM thruster integrated with the gimbal; conditions were 

10.0 Grms, through each of the 3 axes for 2-min duration. The 

thruster performance was nominal pre- and post-vibration. 

Thermal vacuum testing of the PM thruster was conducted 

using flight-approved methods and the Dawn spacecraft 

environmental requirements and test procedures. Test 

parameters included -120 °C cold and 203 °C hot limits, and 3 

cycles with hot and cold dwells, with hot and cold thruster 

starts. Testing was truncated during the 3rd cold cycle due to 

minor damage incurred during prior vibration testing. 

Functional testing of the thruster was subsequently completed 

with all data nominal. The PM thruster is presently undergoing 

minor rework at Aerojet in preparation for system integration 

testing at GRC and a second round of qualification-level 

environmental testing at JPL. The thruster will then be 

incorporated into the life validation program upon completion 

of testing. 

 

2

The PM thruster incorporates flight-level design and fabrication 

processes. The thruster mass is 12.7 kg (13.5 kg with cable harness). 

NEXT Thruster String 

PPU 

LPA 

Gimbal 

Thruster 

DCIU 

HPA 
NASA/TM—2008-214988 3

Manufacturing of components for a second PM thruster 

have been completed at Aerojet and are available for assembly 

in support of 1st-user needs; e.g., qualification unit thruster. 

Thruster documentation (work instructions and drawings) will 

be formally released under the contractor configuration 

management process prior to assembly of the second PM 

thruster. 

B. Power Processor Unit 

Manufacturing of an engineering model (EM) PPU is 

nearing completion at L3 Comm ETI (L3) (fig. 4).3

 The PPU 

is planned to be delivered to GRC for thruster integration 

testing and then qualification-level vibration and thermal 

vacuum testing back at L3. The PPU will subsequently be 

used in systems integration testing which will include 

EMI/EMC measurements. 

The PPU incorporates a modular beam power supply and 

improved packaging that provides performance and produceability benefits over the NSTAR PPU approach. It has a 

flexible, scalable architecture which can be adapted to a wide 

range of thrusters4

 and missions, with a wide throttle range 

capability (250 to 7200 W), a > 0.2 kW/kg specific power, and 

a simple thermal interface (a 65 °C baseplate). 

Additional design elements of the PPU include the following features described below. The PPU has modular construction designed for manufacturability and low cost. The input is 

designed for a wide range unregulated input bus (80 to 

160 Vdc), with separate 100 and 28 V input filters. The 

balanced filters conform to JPL requirements (isolated return). 

The PPU has separate thruster outputs capable of operating 

two thrusters, one at a time. It also is designed to accommodate independent neutralizer keeper operation so that the PPU 

can operate the beam of one thruster combined with the 

neutralizer from a second thruster. The PPU has a built-in grid 

clear circuit with variable programmable duration and current 

capability up to the full output current of the discharge power 

supply. 

C. Propellant Management System 

The EM PMS (fig. 5) delivers low pressure gas to the thruster 

from a supercritical xenon supply source, and it consists of a 

HPA and LPA. The PMS provides independent xenon flow 

control to the thruster main discharge, and discharge and 

neutralizer cathodes. Xenon flow control is achieved by use of a 

thermal throttle flow control device (FCD) and a proportional 

flow control valve (PFCV). The FCD is a porous metal plug 

which provides a desired flow rate for a given inlet pressure and 

 

3

The PPU design was fully documented prior to manufacturing with 

all drawings and assembly instructions in place and under configuration management prior to fabrication. 4

The NEXT PPU is fully-compatible with operating the NSTAR 

thruster over its full throttle table; this was successfully demonstrated 

using the NEXT breadboard PPU on an EM NSTAR thruster. 

operating temperature. Fine control of flow rate is achieved by 

precise, active control of the FCD inlet pressure while the 

operating temperature of the porous metal plug is controlled at 

a constant setpoint. The inlet pressure and operating temperature of the FCD are actively controlled by DCIU simulator. 

The inlet pressure is controlled using the PFCV with pressure 

transducer signal feedback while the operating temperature is 

controlled using a resistive element heater and temperature 

sensor feedback signal. This NEXT PMS design concept 

(ref. 13) provides significant mass and volume reductions in 

the system as compared to the bang-bang regulation scheme 

implemented for the Deep Space 1 or Dawn missions 

(ref. 14).5

 

Figure 4.—EM Power Processor Unit at L3. 

Figure 5.—Schematic of EM Propellant Management System. 

 

5

The NEXT PMS is fully-compatible with the NSTAR thruster, 

operable over its entire throttling table with greater flow precision 

and independent flow control of all 3 feeds to the thruster. 

L

T

P

KEY:

P

FCD Inlet Filter 

Low

Pressure 

Assembly T

P P

T

Main 

Flow 

Cathode 

Flow 

Neutralizer 

Flow

T T T

L

P

L

P T P P

 Xe 

T

P

P

High 

Pressure 

Assembly P P P

T T

PFCV with Filter 

Service Valve 

Latching Valve 

Pressure Transducer 

with Internal Platinum 

RTD

Thermal Throttle

(Redundant 

Components) 
NASA/TM—2008-214988 4

Aerojet has completed manufacturing of the EM PMS 

elements, including 2 HPAs (one flight-like) and 3 LPAs (oneflight like).6

 All assemblies have completed functional testing, 

and both flight-like HPA and LPA assemblies have successfully completed qualification-level vibration testing 

(14.10 Grms through each of the 3 axes for 2 min duration) 

(ref. 12) and thermal vacuum testing (four thermal cycles from 

12 to 70 °C). The assemblies have been delivered to GRC and 

are being prepared for system integration testing. 

D. Digital Control Interface Unit 

The DCIU is the primary interface between the spacecraft 

and the ion propulsion system. It performs all the command and 

control functions for the PMS and the PPU. Under the scope of 

the NEXT program only DCIU simulators have been developed. The DCIU simulator consists of a computer, test support 

equipment, EM PMS pressure loop control cards, and the 

associated algorithms to control the PMS and PPU and receive 

telemetry. The DCIU simulators are capable of operating a 3 

thruster string system, and are being used to validate control 

algorithms, support PPU input/output testing and single-string 

and multi-string integration tests, and operate a PMS kernel 

during thruster life testing. Fully integrating the DCIU functionality into the PPU is under consideration as an approach that 

may be implemented in a follow-on PPU build cycle. Elimination of the separate DCIU should reduce system cost and 

complexity. 

E. Gimbal 

JPL has completed the development of a breadboard gimbal 

for the NEXT thruster (ref. 15). Designed and fabricated by 

Swales Aerospace, the gimbal is of a flight-like design using 

JPL-approved materials with certifications. The mass of the 

gimbal is less than 6 kg (lighter than the Dawn gimbal used 

for the smaller NSTAR thruster), and has a two-axis range of 

motion: ±19°, ±17°. The gimbal has successful completed 

integration and functional testing with the PM thruster (fig. 6), 

and has passed two qualification level vibration tests and lowlevel shock tests, with minor issues. Few if any modifications 

are anticipated to transition to flight. However torque margin 

tests with the thruster harness and propellant line routing 

attached will be required. 

F. System Integration 

Single string integration testing was conducted in Phase 1 

(ref. 10). This test included an EM thruster with a breadboard 

PPU and breadboard PMS. 

Early in Phase 2 a multi-thruster array test (MTAT) was 

executed at GRC, focusing on the characterization of individual thruster, and array, performance and behavior – as affected 

 

6

The non-flight like assemblies are identical except for the use of 

lower cost equivalent parts.

Figure 6.—Gimbal Assembly and PM Thruster 

undergoing Articulation Testing at JPL. 

Figure 7.—Three NEXT EM Thrusters at full power in 

Multi-Thruster Array Test at NASA GRC. 

by the simultaneous operation of multiple ion thrusters 

(ref. 12). The subject of this effort was a four engineering 

model NEXT thruster array in a 3+1 flight-representative 

configuration where one thruster was dormant (a spare) (fig. 7). 

This test was executed concurrent with detailed plasma 

environments and plume measurements (refs. 16 to 19). The 

array was operated over a broad range of conditions including 

the simultaneous firing of 3 thrusters at 20.6 kW total input 

power, yielding a total thrust of about 710 mN, at 4190 s 

specific impulse and approximately 71% efficiency. Major 

findings from the MTAT include: the performance observed 

for a thruster during operation in an array configuration 
NASA/TM—2008-214988 5

appears to be consistent with that measured during singular 

thruster operation with no apparent deleterious interactions; and, 

operation of 1 neutralizer to neutralize 2-or-more thruster 

beams appears to be a potentially viable fault-recovery mode, 

and viable system architecture with significant system 

performance advantages. Overall, the results indicated single 

thruster operations are generally independent of array 

configuration. This finding may have significant implications 

with respect to testing requirements and architectural flexibility for multi-thruster systems. 

Single-and 3-thruster string system integration tests are 

planned during Phase 2, to include operation of the PM thruster 

with the EM PPU and EM PMS (all under vacuum), 

with/without the addition of other thrusters into the system, all 

under command of the DCIU simulator. An exhaustive series of 

tests are planned, the scope of which is to verify that the 

integrated system of NEXT components meets the project 

requirements and verify the interfaces between the system 

components. Primary objectives of the single-string7

 test 

include: demonstrating operation of thruster over the full NEXT 

throttle table with the EM PPU and EM PMS; demonstrating 

operation of the system at off-nominal conditions; and demonstrating recycle and fault protection operation. In total over 100 

system level requirements will be validated. Primary objectives 

of the multi-thruster string test include: validating DCIU and 

PMS functionality supporting multi-thruster operation; and 

documenting any potential subsystem interactions. 

The multi-thruster system hardware completed in Phase 2 

will remain as a flight-system-test bed (representative of 

expected flight propulsion system) available for continued use 

as a risk reduction utility to support 1st- user needs. 

G. Life Validation 

Validation of the NEXT thruster life requirement is being 

addressed via a combination of test and analyses.8

 An EM 

thruster successfully completed a 2000 hr wear test (ref. 9). A 

second EM thruster (with PM ion optics manufactured by 

Aerojet) is presently undergoing long duration life testing at 

GRC at full power having accumulated approximately 9300 hr 

of operation to date, processing over 191 kg of xenon. The 

thruster has demonstrated over 7.9 × 106

 N-s total impulse; the 

highest total impulse ever demonstrated by an ion thruster in 

the history of space propulsion. The EM thruster performance 

and wear rates of critical thruster components are consistent 

with model predictions. The life test is scheduled to continue 

at full power until the 300 kg requirement is demonstrated at 

which point the thruster will be power throttled to a mission 

profile and continue operation. The life capability of the PM 

thruster will be established by a combination of (a) results 

 

7

Single-string refers to a single Thruster – PPU – LPA PMS 

configuration. 

8

The 300 kg xenon propellant throughput requirement corresponds to 

about 14.5 khr of operation at full power; 21.7 khr for qualificationlevel testing. 

from the on-going EM thruster life testing, (b) similarity 

analysis to the EM thruster, (c) thruster service life modeling, 

and (d) a full-power wear test of a PM thruster. 

A comprehensive service life assessment of the NEXT 

thruster has been conducted at GRC (ref. 20). The assessment 

involved the application of several models to evaluate all the 

known failure modes. The assessment conservatively predicts 

that the earliest failure would be the accelerator grid, with a 

structural failure occurring sometime after 730 kg of xenon 

throughput. At full power, this would be over 35 khr of 

operation. Other failure modes are predicted to occur at a 

condition in excess of 800 kg thruster xenon throughput; well 

beyond the mission-derived life time requirement of 300 kg 

xenon throughput. The analyses will be refined as results from 

thruster extended duration tests are available. 

III. System Description 

Current best estimates for NEXT flight subsystem characteristics (mass and volume) are listed in Table II. 

TABLE II.—SUBSYSTEM CHARACTERISTICS 

Resource System Current Best Estimate Basis 

Thruster 12.7 PM Actual, w/o 

harness 

PPU 34.5 EM Actual 

PMS HPA 1.9 EM Actual, w/plate & 

Test Support Equip. 

PMS LPA 3.1 EM Actual, w/plate & 

Test Support Equip. 

Mass 

(kg) 

Gimbal 6 Breadboard Actual 

Thruster 55 dia by 44 length PM Actual 

PPU 42 by 53 by 14 EM Actual

PMS HPA 33 by 15 by 6.4 EM Actual

PMS LPA 38 by 30.5 by 6.4 EM Actual

Envelope 

(cm) 

Gimbal 72 cm corner-corner, 

61 cm flat-flat Breadboard Actual 

The NEXT ion propulsion system (IPS) is anticipated to 

consist of 1 or more thruster strings (see fig. 1). A thruster 

string is composed of a thruster/gimbal assembly, a PPU, and 

a PMS LPA. An HPA and a DCIU9

 would complete the 

system. Thruster strings are added for ‘mission performance 

reasons (to process the available power) and an additional 

string is added for failure tolerance (a spare). A possible 

NEXT IPS 3-string configuration is illustrated in figure 8. 

There are a number of system design characteristics that 

would likely be implemented in a manner different than the 

Dawn IPS: 

 

9

Alternately, all DCIU functionality could be incorporated into the 

PPU. This is the intended design approach for a 2nd generation 

NEXT PPU. 
NASA/TM—2008-214988 6

PPU|DCIU

A

LPA

A

Xe TANK

HPA

SPACECRAFT 28

V POWER

100 V POWER

CMD & TLM

Xe

28 V POWER

CONDITIONED LV POWER

PPU|DCIU

B

LPA

B

PPU|DCIU

C

LPA

C

G

I

M

B

A

L

NEXT ION PROPULSION SYSTEM

SOLAR ARRAY

GIMBAL ACTUATION

CMD & TLM

SPACECRAFT

CONTROLLER

G

I

M

B

A

L

THRUSTER

G

I

M

B

A

L

THRUSTER

THRUSTER

Figure 8.—NEXT IPS 3-String ('2+1') Configuration. 

9 Although the EM NEXT PPU was designed to operate 

2 thrusters, the Dawn IPS issues with relays that switch 

the power supply outputs between thrusters and the associated software development requirements indicate 

this to be an unattractive approach with negative system-level reliability consequences. Therefore, the 

NEXT IPS illustrated in figure 8 assumes no crossstrapping implementation, with direct wiring from each 

PPU to 1 corresponding thruster. 

9 The DS1 and Dawn IPS had separate DCIU boxes, one 

consequence of which is increased system cost and 

complexity. Much of the functionality normally associated with a separate DCIU as implemented in DS1 and 

Dawn has already been propagated into the NEXT EM 

PPU (thruster sequences). The project intends on building a 2nd-generation PPU which propagates the remaining DCIU functionality (PMS controls) into the PPU 

box. Figure 8 illustrates this approach, where each PPU 

and LPA is controlled by a separate digital interface contained within the PPU box. Each of these dedicated 

DCIUs is capable of commanding the single HPA. 

9 The gimbal drive electronics are assumed to be maintained as a spacecraft function and are not processed 

through the DCIU as implemented on Dawn. 

Spacecraft integration should be simplified with NEXT due 

to improved capabilities and features: less thruster strings per 

mission total impulse; modular, simplified xenon feed system; 

the PPU is compatible with wider baseplate thermal range than 

NSTAR; and the gimbal has smaller footprint than NSTAR. 

IV. Performance 

A. System 

NEXT provides significant performance benefits as compared to NSTAR SOA. The higher propellant throughput 

capability of the thruster and its greater total impulse allow it 

to accomplish low power, high ∆V missions (Dawn-like) with 

fewer numbers of thrusters. The higher power and thrust 

capability of NEXT allow power-driven missions (outerplanet) to be accomplished with fewer thruster strings. The 

higher specific impulse capability of the NEXT thruster 

reduces the spacecraft propellant mass, thus accommodating 

more payload mass. The increased system power throttling 

range (11.9-to-1 PPU input power range for NEXT versus 4.8-

to-1 for NSTAR) allows single thruster string operation over a 

much broader range of solar distances. Table III lists throttle 

table performance data for the NEXT thruster (representative 
NASA/TM—2008-214988 7

of both EM and PM thrusters) and data obtained from the 

breadboard PPU. 

TABLE III.—DEMONSTRATED NEXT 

THRUSTER AND PPU PERFORMANCE 

Thrust, 

mN 

Isp

a

, 

s ηThrusterb PPU - 

Pin, kW ηPPUc ηSystemd

Xe Mass Flow Rate = 5.76 mg/s 

236 4190 0.71 7.220 0.95 0.67 

221 3910 0.70 6.385 0.95 0.66 

208 3690 0.69 5.780 0.95 0.65 

192 3395 0.68 4.965 0.95 0.64 

Xe Mass Flow Rate = 5.12 mg/s 

208 4150 0.70 6.390 0.95 0.66 

194 3875 0.69 5.660 0.95 0.65 

184 3660 0.68 5.085 0.95 0.65 

169 3360 0.67 4.490 0.95 0.63 

Xe Mass Flow Rate = 4.46 mg/s 

181 4150 0.70 5.600 0.95 0.66 

169 3875 0.69 4.920 0.95 0.65 

160 3660 0.68 4.455 0.95 0.64 

147 3360 0.67 3.860 0.94 0.63 

137 3125 0.65 3.425 0.94 0.61 

Xe Mass Flow Rate = 3.92 mg/s 

158 4100 0.69 4.870 0.95 0.65 

147 3830 0.68 4.315 0.95 0.64 

139 3615 0.67 3.910 0.94 0.63 

128 3325 0.66 3.390 0.94 0.62 

119 3090 0.64 3.010 0.94 0.60 

Xe Mass Flow Rate = 3.16 mg/s 

134 4310 0.71 4.235 0.95 0.67 

125 4025 0.70 3.760 0.94 0.66 

118 3800 0.69 3.415 0.94 0.64 

108 3490 0.67 2.970 0.94 0.62 

101 3250 0.65 2.620 0.94 0.61 

Mass Flow Rate = 2.60 mg/s 

107 4190 0.68 3.460 0.94 0.64 

99.9 3910 0.67 3.080 0.94 0.62 

94.3 3690 0.65 2.765 0.94 0.61 

86.7 3395 0.64 2.415 0.94 0.60 

80.6 3155 0.62 2.160 0.94 0.58 

Xe Mass Flow Rate = 2.05 mg/s 

80.2 4000 0.65 2.585 0.94 0.61 

74.9 3735 0.63 2.300 0.94 0.60 

70.7 3525 0.62 2.090 0.94 0.59 

65.0 3240 0.61 1.825 0.93 0.57 

60.4 3015 0.59 1.635 0.93 0.55 

57.8 2885 0.58 1.520 0.94 0.54 

55.1 2745 0.56 1.415 0.94 0.53 

49.2 2450 0.53 1.210 0.93 0.49 

48.1 2400 0.52 1.175 0.93 0.48 

37.2 1855 0.43 0.865 0.91 0.39 

31.8 1585 0.37 0.740 0.90 0.33 

Xe Mass Flow Rate = 1.85 mg/s 

25.5 1400 0.33 0.610 0.89 0.29 a

Isp = specific impulse b

ηThruster = thruster efficiency c

ηPPU = PPU efficiency obtained at 100 V input bus voltage d

ηSystem= system efficiency = ηThruster ● ηPPU

System efficiencies, defined here as the product of the 

thruster and PPU efficiencies, are also listed in the throttle 

table. As noted from the NEXT throttle table, the PPU input 

power ranges from 7.22 kW down to 0.61 kW, corresponding 

to a specific impulse range of about 4190 to 1400 s, and thrust 

range from about 236 mN down to about 26 mN. Over this 

input power range the thruster, PPU, and system efficiencies 

vary from about 0.71-0.33, 0.95-0.89, and 0.67-0.29 respectively. 

B. Mission 

The NEXT project began with Flagship-class Deep Space 

Design Reference Missions (a Titan Explorer, and Neptune 

Orbiter, both assuming aerocapture at the destinations) as the 

‘design driver’ mission applications. A refocus study was 

conducted in 2004 to investigate NEXT IPS applications to 

both Discovery- and New Frontiers-class missions. Several 

Discovery-class mission studies shows that NEXT outperforms SOA NSTAR, yielding higher net payload mass with 

fewer thrusters. NEXT also was either mission-enhancing 

(relative to chemical propulsion) or mission-enabling for 

several New Frontiers- and Flagship-class studies. A summary 

of these findings are listed in table IV which indicates that 

NEXT provides mission benefits across all planetary science 

mission classes. First-user costs however are identified as an 

area of concern. 

TABLE IV.—SUMMARY OF NEXT IPS MISSION BENEFITS 

Mission NEXT Mission Benefits 

Discovery- Small Body Missions: 

9 Near Earth Asteroid Rendezvous 

9 Vesta-Ceres Rendezvous 

(Dawn) 

9 Comet Rendezvous 

9 Deimos Sample Return 

Higher net payload mass with fewer 

thrusters than NSTAR system 

New Frontiers: 

9 Comet Surface Sample Return 

Higher net payload mass than 

NSTAR, with, simpler EP System: 

2+1 NEXT versus 4+1 NSTAR 

thrusters 

New Frontiers: 

9 Titan Direct Lander 

>700 kg entry package with 1+1 

NEXT system, potentially within 

New Frontiers cost cap 

Flagship - Saturn System Missions: 

9 Titan 

9 Enceladus 

> 2400 kg to Saturn Orbit Insertion 

with 1+1 NEXT system, EGA and 

Atlas 5 EELV - doubles delivered 

mass of chemical/JGA approach 

> 4000 kg to Saturn Orbit Insertion 

with 3+1 NEXT system, EGA and 

Delta IV Heavy 

V. Transition to Flight 

A number of project activities are being conducted to increase the likelihood of transitioning the NEXT IPS technology to flight in the near-term, and do so with minimal 

technical difficulties and at low cost. These activities include: 

reviewing Dawn IPS ‘lessons-learned’ and implementing 

strategies to mitigate the likelihood of experiencing similar 

difficulties; conducting independent reviews of NEXT 

technology status with representation from the user community and incorporating the feedback into the development plan; 
NASA/TM—2008-214988 8

and identifying additional technology development and 

validation activities which may be of value in transitioning the 

TRL6 IPS technology to flight and reduce barriers to 1st-user 

implementation (reduce non-recurring costs, etc.). These 

aspects are discussed further. 

The NEXT project has placed particular emphasis on key 

aspects of IPS development with the intention of avoiding the 

difficulties experienced by the Dawn mission in transitioning 

the NSTAR-based technology to an operational ion propulsion 

system (ref. 7). Some of the issues experienced during Dawn 

IPS development, and mitigation strategies implemented under 

NEXT IPS development are listed in table V. 

TABLE V.—DAWN ISSUES AND NEXT MITIGATIONS 

Element Dawn IPS Issue NEXT IPS Mitigation Strategy 

Inadequate documentation to reproduce 

NSTAR/DS1 thrusters 

9 Proto-flight thruster build phase 

executed 

9 Proto-flight thruster drawings 

and work instructions to be 

released under contractors flight 

hardware configuration management system at completion 

of Phase 2 

Difficulty in assembling and alignment of 

ion optics 

Improved manufacturing processes 

and mounting design reduce 

alignment touch labor and improve 

alignment precision Thruster 

Design changes 

identified by on-going 

thruster life test needed 

to be implemented into 

the flight thruster 

design and build 

process 

9 Design changes identified by 

(early) execution of wear- and 

long-duration tests of NEXT 

thrusters have already been 

incorporated into the PM design

9 Mission-derived requirement of 

300 kg propellant throughput 

will be demonstrated by 1st

-

quarter 2008; well in advance 

of commitment to flight 

Inadequate documentation to reproduce 

NSTAR/DS1 PPUs 

PPU manufactured by contractor’s 

flight production group with all 

documentation - manufacturing 

drawings and assembly instructions 

– completed and under configuration control prior to EM PPU build 

The DS1 PPU was not 

designed to be 

manufacturable 

resulting in fabrication, 

assembly and test issues 

The NEXT PPU has a modular 

construction 

which is: designed for semiautomatic testing at module level; 

designed for manufacturability; 

relatively low cost; and contains 

parts which have short lead time, 

are new, and available 

PPU 

No thermal cycle 

testing or analysis of 

the NSTAR/DS1 PPU 

was completed prior to 

Dawn 

Thermal cycle testing and analysis 

of the NEXT PPU will be 

completed during Phase 2 

PMS 

Dawn PMS (a 3-

thruster version of the 

DS1 system) is 

complex, bulky, and 

required extensive 

modification to satisfy 

mission requirements 

9 DS1 and Dawn feed system 

engineers heavily involved in 

NEXT design from project 

initiation 

9 NEXT PMS design is 

lightweight and compact 

An independent review (IR) of the NEXT project was 

conducted with the objective of providing an assessment of the 

NEXT project and its ability to meet the project objectives. 

Specifically the IR board, which included participation from 

the potential user community, was asked to review the planned 

validation and qualification tests and analyses for the baseline 

project, and identify project risks and provide recommendations for project modifications. With implementation of 11 

findings, the IR board concluded that the NEXT project is on a 

path that would effectively support future users. The NEXT 

team reviewed and concurred with all of the IR findings and is 

in the process of implementing the board findings within the 

constraints of the development program. 

As a product of the Dawn lessons learned, the IR findings, 

and EP architecture studies conducted under ISPT, the NEXT 

project has identified additional technology development and 

validation activities which may facilitate the transition of the 

NEXT IPS to flight and reduce barriers to 1st-user implementation. These tasks (see table VI) include manufacturing of a 

second PM thruster under configuration control which could 

be available as a qualification unit to the 1st-user, fabrication 

of an updated (2nd-generation) NEXT PPU with integral 

DCIU, and use of the NASA Multi-Thruster Array as a flightrepresentative propulsion system test bed. It is anticipated that 

by taking these additional transition steps the 1st-user IPS 

implementation costs will be reduced.10 

TABLE VI.—NEXT IPS TRANSITION TO FLIGHT ACTIVITIES 

System Element Activity 

Thruster 

9 Complete 2nd PM thruster under configuration 

control 

9 Update thruster design and analysis as 

required 

PPU/DCIU 

9 Manufacture 2nd-generation PPU incorporating all DCIU functionality internal to PPU 

9 Complete qualification-level testing of 2ndgeneration PPU 

9 Update PPU analysis as required, including 

reliability 

System 

Maintain NEXT Multi-Thruster Array – with 

integrated EM and PM thrusters, EM PPU, EM 

PMS, and gimbal - as a multi-engine flightsystem-test bed to support flight system 

integration activities 

VI. Summary 

NEXT project activities have brought next-generation ion 

propulsion technology to a mature state, with existing tasks 

completing the majority of the NEXT technology validation. 

The NEXT project - at the completion of CY07 – should be in 

a high state of technical readiness to support FY08 Announcement of Opportunities (AO) relative to Discovery- and 

New Frontier-class missions. Functional and qualificationlevel environmental testing of key system elements are 

 

10A NASA GRC/JPL jointly-developed cost model for the NEXT IPS 

estimates 1st-user costs to be approximately 50% of the Dawn IPS. 
NASA/TM—2008-214988 9

scheduled to be completed; the thruster life test should exceed 

the throughput demonstrated on the NSTAR thruster; and 

system integration testing with the most mature hardware 

products will be drawing to a close. The project has identified 

a number of additional tasks to facilitate transition to first 

flight and these have been recommended to ISPT. 

NEXT IPS technologies will be ready for upcoming competed and directed robotic mission opportunities in the 

FY08/09 timeframe. With additional investments in transitioning the technology to flight, NEXT IPS 1st-user costs should 

be about half that experienced with the Dawn mission IPS. 

References 

1. Benson, S., Patterson, M., Vaughan, D., Wilson, A., and Wong, 

B., “NASA’s Evolutionary Xenon Thruster (NEXT) Phase 2 

Development Status,”AIAA-2005-4070, Joint Propulsion Conference, July 2005. 

2. Patterson, M.J., Foster, J.E., Haag, T.W., Soulas, G.C., Pastel, 

M.R., and Roman, R.F., “NEXT: NASA’S Evolutionary Xenon 

Thruster Development Status,” AIAA-2003-4862, Joint Propulsion Conference, July 2003. 

3. Todd, P. Wiseman, S., Martinelli, R., and Pinero, L., “Status of 

the NEXT 7 kW Power Processing Unit,” AIAA-2005-3868, 

Joint Propulsion Conference, July 2005. 

4. Hoskins, W., Polaha, J., Hobson, D., Soulas, G., Patterson, M., 

Sovey, J., and Talerico, L., “Overview of the NEXT Ion Propulsion System Program at Aerojet,” AIAA-2005-3885, Joint Propulsion Conference, July 2005. 

5. Monheiser, J., Aadland, R., and Wilson, F., “Development of a 

Ground Based Digital Control Interface Unit (DCIU) for the 

NEXT Propulsion System,” AIAA-2004-4112, Joint Propulsion 

Conference, July 2004. 

6. Brophy, J.R., et al., “Ion Propulsion System (NSTAR) DS1 

Technology Validation Report,” JPL Publication 00-10, October 

2000. 

7. Brophy, J.R., “Preliminary Dawn IPS Lessons Learned,” JPL, 

September 9, 2005. 

8. Soulas, G.C., Domonkos, M.T., and Patterson, M.J., “Performance Evaluation of the NEXT Ion Engine, AIAA-2003-5278, 

Joint Propulsion Conference, July 2003. 

9. Soulas, G.C., et al., “NEXT Ion Engine 2000 Hour Wear Test 

Results,” AIAA-2004-3791, Joint Propulsion Conference, July 

2004. 

10. Patterson, M.J., Pinero, L., Aadland, R., and Komm, D., “NEXT 

Ion Propulsion System: Single-String Integration Test Results,” 

JANNAF Conference, May 2004. 

11. Patterson, M.J., et al., “NEXT: NASA’S Evolutionary Xenon 

Thruster Development Status,” AIAA-2003-4862, Joint Propulsion Conference, July 2003. 

12. Patterson, M.J., Foster, J., McEwen, H., Pencil, E., Van Noord, 

J., and Herman, D., "NEXT Multi-Thruster Array Test - Engineering Demonstration," AIAA-2006-5180, Joint Propulsion 

Conference, July 2006. 

13. Aadland, R.S., Frederick, H., Benson, S.W., and Malone, S., 

“Development Results of the NEXT Propellant Management 

System,” JANNAF Conference, 2005. 

14. Brophy, J.R, et al., “Status of the Dawn Ion Propulsion System,” 

AIAA-04-3433, Joint Propulsion Conference, July 2004. 

15. Snyder, J.S, O’Connell, M.R., Fernandez, J.P., Wang, G., 

McNabb, R.S., and Crumb, D., “Vibration Test of a Breadboard 

Gimbal for the NEXT Ion Engine,” AIAA Paper No. 2006-4665, 

Joint Propulsion Conference, July 2006. 

16. Pencil, E.J., Foster, J.E., Patterson, M.J., Diaz, E., Van Noord, 

J., and McEwen, H., "Ion Beam Characterization of NEXT 

Multi-Thruster Array Plume," AIAA-2006-5182, Joint Propulsion Conference, July 2006. 

17. Foster, J.E., Pencil, E., Patterson, M.J., McEwen, H., Diaz, E., 

and Van Noord, J., "Plasma Characteristics Measured in the 

Plume of a NEXT Multi-Thruster Array," AIAA-2006-5181, 

Joint Propulsion Conference, July 2006. 

18. McEwen, H., Foster, J.E., Pencil, E., Patterson, M.J., Diaz, E., 

and Van Noord, J., "Characterization of Plasma Flux Incident on 

a Multi-Thruster Array," AIAA-2006-5183, Joint Propulsion 

Conference, July 2006. 

19. Foster, J.E., Pencil, E., McEwen, H., Patterson, M.J., Diaz, E., 

and Van Noord, J., "Neutralizer Plasma Coupling in a NEXT 

Multi-thruster array", AIAA-2006-5184, Joint Propulsion Conference, July 2006. 

20. Van Noord, J., “NASA’s Evolutionary Xenon Thruster (NEXT) 

Ion Propulsion System - Technology for NASA Space Science 

Missions - Service Life Assessment,” GRC-NEXT-103, December 21, 2005. 
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NEXT Ion Propulsion System Development Status and Capabilities 

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Patterson, Michael, J.; Benson, Scott, W. 

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14. ABSTRACT

NASA’s Evolutionary Xenon Thruster (NEXT) project is developing next generation ion propulsion technologies to provide future NASA 

science missions with enhanced mission performance benefit at a low total development cost. The objective of the NEXT project is to 

advance next generation ion propulsion technology by producing engineering model system components, validating these through 

qualification-level and integrated system testing, and ensuring preparedness for transitioning to flight system development. As NASA’s 

Evolutionary Xenon Thruster technology program completes advanced development activities, it is advantageous to review the existing 

technology capabilities of the system under development. This paper describes the NEXT ion propulsion system development status, 

characteristics and performance. A review of mission analyses results conducted to date using the NEXT system is also provided.

15. SUBJECT TERMS

Propulsion; Electric propulsion; Electrostatic propulsion; Ion propulsion; Ion optics 

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