Michael J. Patterson and Scott W. Benson

Glenn Research Center, Cleveland, Ohio

NEXT Ion Propulsion System Development Status

and Performance

NASA/TM—2008-214986

January 2008

AIAA–2007–5199
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Michael J. Patterson and Scott W. Benson

Glenn Research Center, Cleveland, Ohio

NEXT Ion Propulsion System Development Status

and Performance

NASA/TM—2008-214986

January 2008

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Prepared for the

43rd Joint Propulsion Conference

sponsored by the American Institute of Aeronautics and Astronautics

Cincinnati, Ohio, July 8–11, 2007

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Level of Review: This material has been technically reviewed by technical management.
NASA/TM—2008-214986 1

NEXT Ion Propulsion System Development Status and Performance 

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 and prototype model system components, 

validating these through qualification-level and integrated system testing, and ensuring preparedness for 

transitioning to flight system 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. 

Nomenclature 

BOL Beginning-of-Life 

DCIU Digital Control Interface Unit 

EGA Earth Gravity Assist 

EM Engineering Model 

F Thrust, mN 

FCD Flow Control Device 

GRC Glenn Research Center 

HPA High Pressure Assembly 

JPL Jet Propulsion Laboratory 

I Thruster total impulse, Newton seconds 

Ib Beam current, Amperes 

IPS Ion Propulsion System 

Isp Specific impulse, s 

JPA Jupiter Gravity Assist 

LPA Low Pressure Assembly 

Lt Projected thruster life time, khr 

MT Mass flow rate into thruster, mg/s 

MT-TOT Thruster total propellant throughput capability, kg of Xenon 

NEXT NASA’s Evolutionary Xenon Thruster project 

NSTAR NASA Solar Electric Propulsion Technology Application Readiness 

PFCV Proportional Flow Control Valve 

Pin-PPU PPU input power, kW (unless otherwise specified) 

Pin-t Thruster input power, kW (unless otherwise specified) 

PM Prototype Model 

PMS Propellant Management System 

PPU Power Processor Unit 

TRL Technology Readiness Level 

Vbps Beam power supply voltage, Volts 

ηPPU Demonstrated PPU efficiency (breadboard, at 100 V input bus voltage) 

ηt Demonstrated Thruster efficiency (PM) 

ηSystem System efficiency, product of thruster and PPU efficiencies 
NASA/TM—2008-214986 2

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 competitivelyselected NASA Research Announcement 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, high-efficiency 7-kW power processor unit (PPU); a highly flexible advanced xenon propellant 

management system (PMS) 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 SOA NSTAR performance 

characteristics are listed in table 1. 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 1.—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% 

 

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. 

NEXT Thruster String 

PPU 

LPA 

Gimbal 

Thruster 

DCIU 

Functionality 

HPA 

Figure 1.—NEXT Ion Propulsion System Elements.
NASA/TM—2008-214986 3

Figure 2.—NEXT makes high-performing ion propulsion systems available to a wide range of 

NASA missions. 

NEXT technology is applicable to a wide range of NASA solar system exploration missions 

(see fig. 2), as well as earth-space 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. 

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. 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 project requirement is >300 kg (1.23×107

 N-s total impulse), 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 2 000 hr wear test at full power (ref. 9), integration testing with a breadboard PPU 

and breadboard PMS (ref. 10) and structural 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 

9200 hr of operation at full power. 

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. The thermal development test proceeding environmental tests was performed to develop 

 

2

The PM thruster incorporates flight-level design and fabrication processes. The thruster mass is 12.7 kg (13.5 kg with cable 

harness).
NASA/TM—2008-214986 4

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 

(see fig. 3). 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. 

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

 This was preceded by fabrication and 

successful integration test of breadboard PPU during 

Phase 1. The PPU is planned to be delivered to GRC 

for thruster integration testing and then qualificationlevel 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 produce-ability 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). 

 

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. 

Figure 4.—EM Power Processor Unit. 

Figure 3.—PM Thruster undergoing Thermal 

Vacuum Testing at JPL. 
NASA/TM—2008-214986 5

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 (ref. 3). 

C. Propellant Management System 

The EM Propellant Management System (PMS) (fig. 5) delivers low pressure gas to the thruster from 

a supercritical xenon supply source, and it consists of HPA and LPA (ref. 14). 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 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 the 

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 and Dawn missions (ref. 14).5

Figure 5.—Schematic of NEXT 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—2008-214986 6

Aerojet has completed manufacturing of the EM PMS elements, including 2 HPAs (one flight-like) 

and 3 LPAs (one-flight 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. 13) 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 Digital Control Interface Unit (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. 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 singlestring 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 successfully 

completed integration and functional testing with the PM thruster (fig. 6), and has passed two 

qualification level vibration tests and low-level 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 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) (see 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 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 

 

6

The non-flight like assemblies are identical except for the use of lower cost equivalent parts. 
NASA/TM—2008-214986 7

Figure 6.—Gimbal Assembly and PM 

Thruster undergoing Articulation 

Testing at JPL. 

Figure 7.—NEXT EM Thrusters operating at full power in 3+1 

Multi-Thruster Array Test bed at NASA GRC. 

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 singlestring7

 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. 

 

7

Single-string refers to a single Thruster—PPU—LPA PMS configuration. 
NASA/TM—2008-214986 8

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 9200 hr of operation to date, processing over 182 kg of 

xenon. The thruster has demonstrated over 7.8×106 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 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 service life assessment of the NEXT thruster was conducted at GRC for a number of throttle 

conditions (ref. 20). The assessment involved the application of several models to evaluate all the known 

failure modes. For the conditions investigated, the assessment conservatively predicted 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 were 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. Additional analyses of thruster life time over the range of 

the NEXT full-throttle table were conducted, the results of which are discussed in a following section. 

III. Ion Propulsion System Architecture 

Current best estimates for NEXT flight subsystem characteristics (mass, volume, and power) are 

listed in table 2. 

TABLE 2.—NEXT ION PROPULSION SUBSYTEM 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 and Test Support Equipment 

LPA 3.1 EM Actual, w/plate and Test Support Equipment 

Mass 

(kg) 

Gimbal 6 Breadboard Actual 

Thruster 55 dia by 44 length PM Actual 

PPU 42 by 53 by 14 EM Actual

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 

Thruster9 540–6860 PM Actual 

PPU 610–7220 Breadboard Actual 

HPA 4.3 EM Actual PMS LPA 15.9 EM Actual 

Power 

(W) 

Gimbal N/A ------------------------------------------------------- 

 

8

The 300 kg xenon propellant throughput requirement corresponds to about 14.5 khr of operation at full power; 21.7 khr for 

qualification-level testing. 

9

Input power to the thruster from the PPU. 
NASA/TM—2008-214986 9

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. 

The NEXT IPS is anticipated to consist of a series of independent single-strings (see fig. 1). A 

thruster string is composed of a thruster and gimbal assembly, a PPU, and a PMS LPA. An HPA and a 

DCIU10 would complete the system. Thruster strings are added for mission performance reasons (to 

process the available solar 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 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 cross-strapping 

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 

 

10Alternately, all DCIU functionality could be incorporated into the PPU. This is the intended design approach for a 2nd 

generation NEXT PPU. 
NASA/TM—2008-214986 10

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 the NEXT IPS, as compared to NSTAR-based IPS, 

due to improved capabilities and features: 

9 Less thruster strings per mission total impulse are required due to the higher total impulse capability of 

the NEXT thruster; 

9 Modular, simplified xenon feed system; 

9 The NEXT PPU is compatible with wider baseplate thermal range than NSTAR; and 

9 The gimbal has smaller footprint than the NSTAR thruster Dawn gimbal. 

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 

ΔV missions (Dawn-like) with fewer numbers of thrusters. The higher power and thrust capability of 

NEXT allow power-driven missions (outer-planet) 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 3, the NEXT throttle table, documents the performance data for the thruster (representative of 

both EM and PM thrusters) and data obtained from the breadboard PPU. 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 to 

0.33, 0.95 to 0.89, and 0.67 to 0.29 respectively. Figures 9, 10, and 11 show plots of: the thruster and 

system efficiencies vs. specific impulse; the thrust as a function of PPU input power; and the specific 

impulse as a function of PPU input power. 

Analyses of thruster life time were conducted by the author over the entirety of the performance 

conditions of the NEXT throttle table, and these results are also listed in table 3. The calculation 

methodology is described in reference 19, and includes ion optics electrode wear, as well as cathode 

emitter and cathode keeper wear. Thruster life, Lt, in khr, and total Xe propellant throughput capability, 

MT-TOT, in kg, are listed, as well as the corresponding total impulse capability of the thruster, I, in N-s. 

Consistent with prior calculations, the total propellant throughput capability is projected to be in 

excess of >700 kg at full power and over most of the throttle table. However, for the 3 lowest-power 

conditions in the throttle table,11 the throughput capability is projected to be less than the qualification 

requirement of 450 kg. This is due to operation at high magnitudes of accelerator electrode voltage to 

yield high beam current. However, even at these conditions, the projected life time is in excess of 35 khr. 

 

11These 3 throttle conditions were added to the NEXT throttle table after Phase 1 to accommodate investigating NEXT IPS 

applications to Discovery- and New Frontiers-class missions. 
NASA/TM—2008-214986 11

TABLE 3.—NEXT THROTTLE TABLE: DEMONSTRATED THRUSTER AND PPU PERFORMANCE (BOL) 

Vbps, 

V

F, 

mN 

Isp, 

s ηt

Pin-PPU, 

kW ηPPU ηSystem

Lt

khr 

MT-TOT, 

kg Xe I, ×107

 N-s 

MT = 5.76 mg/s; Ib = 3.52 A 

1800 236 4190 0.71 7.220 0.95 0.67 37.6 780 3.20 

1567 221 3910 0.70 6.385 0.95 0.66 37.6 780 2.99 

1396 208 3690 0.69 5.780 0.95 0.65 37.6 780 2.82 

1179 192 3395 0.68 4.965 0.95 0.64 40.5 840 2.80 

MT = 5.12 mg/s; Ib = 3.10 A 

1800 208 4150 0.70 6.390 0.95 0.66 43.0 795 3.22 

1567 194 3875 0.69 5.660 0.95 0.65 43.0 795 3.01 

1396 184 3660 0.68 5.085 0.95 0.65 43.0 795 2.84 

1179 169 3360 0.67 4.490 0.95 0.63 46.3 855 2.81 

MT = 4.46 mg/s; Ib = 2.70 A 

1800 181 4150 0.70 5.600 0.95 0.66 49.1 790 3.20 

1567 169 3875 0.69 4.920 0.95 0.65 49.1 790 2.99 

1396 160 3660 0.68 4.455 0.95 0.64 49.1 790 2.83 

1179 147 3360 0.67 3.860 0.94 0.63 52.9 850 2.80 

1021 137 3125 0.65 3.425 0.94 0.61 65.7 1055 3.23 

MT = 3.92 mg/s; Ib = 2.35 A 

1800 158 4100 0.69 4.870 0.95 0.65 56.5 795 3.21 

1567 147 3830 0.68 4.315 0.95 0.64 56.5 795 2.99 

1396 139 3615 0.67 3.910 0.94 0.63 56.5 795 2.83 

1179 128 3325 0.66 3.390 0.94 0.62 60.8 860 2.80 

1021 119 3090 0.64 3.010 0.94 0.60 75.5 1065 3.23 

MT = 3.16 mg/s; Ib = 2.00 A 

1800 134 4310 0.71 4.235 0.95 0.67 66.4 755 3.20 

1567 125 4025 0.70 3.760 0.94 0.66 66.4 755 2.98 

1396 118 3800 0.69 3.415 0.94 0.64 66.4 755 2.82 

1179 108 3490 0.67 2.970 0.94 0.62 71.5 815 2.79 

1021 101 3250 0.65 2.620 0.94 0.61 88.9 1010 3.22 

MT = 2.60 mg/s; Ib = 1.60 A 

1800 107 4190 0.68 3.460 0.94 0.64 83.6 785 3.22 

1567 99.9 3910 0.67 3.080 0.94 0.62 83.6 785 3.01 

1396 94.3 3690 0.65 2.765 0.94 0.61 83.6 785 2.84 

1179 86.7 3395 0.64 2.415 0.94 0.60 90.1 845 2.81 

1021 80.6 3155 0.62 2.160 0.94 0.58 112 1050 3.25 

MT = 2.05 mg/s; Ib = 1.20 A 

1800 80.2 4000 0.65 2.585 0.94 0.61 110 815 3.18 

1567 74.9 3735 0.63 2.300 0.94 0.60 110 815 2.97 

1396 70.7 3525 0.62 2.090 0.94 0.59 110 815 2.80 

1179 65.0 3240 0.61 1.825 0.93 0.57 119 875 2.78 

1021 60.4 3015 0.59 1.635 0.93 0.55 147 1090 3.21 

936 57.8 2885 0.58 1.520 0.94 0.54 196 1445 4.07 

850 55.1 2745 0.56 1.415 0.94 0.53 293 2160 5.80 

679 49.2 2450 0.53 1.210 0.93 0.49 366 2705 6.49 

650 48.1 2400 0.52 1.175 0.93 0.48 212 1565 3.68 

400 37.2 1855 0.43 0.865 0.91 0.39 49.3 365 0.66 

300 31.8 1585 0.37 0.740 0.90 0.33 36.5 270 0.42 

MT = 1.85 mg/s; Ib = 1.00 A 

275 25.5 1400 0.33 0.610 0.89 0.29 46.3 310 0.42 
NASA/TM—2008-214986 12

0.200

0.300

0.400

0.500

0.600

0.700

0.800

1000 1500 2000 2500 3000 3500 4000 4500

Specific Impulse, seconds

Efficiency

Thruster

System

Figure 9.—NEXT Thruster and System Efficiency versus Specific Impulse. 

0

50

100

150

200

250

0 1000 2000 3000 4000 5000 6000 7000 8000

PPU Input Power, W

Thrust, mN

Figure 10.—NEXT Thrust as a function of PPU Input Power. 
NASA/TM—2008-214986 13

1000

1500

2000

2500

3000

3500

4000

4500

0 1000 2000 3000 4000 5000 6000 7000 8000

PPU Input Power, W

Specific Impulse, seconds

Figure 11.—NEXT Specific Impulse as a function of PPU Input Power. 

100

1,000

10,000

100,000

1,000,000

0 1000 2000 3000 4000 5000 6000 7000 8000

Thruster Input Power, W

Parameter

Life, khr

Throughput, kg Xe

Figure 12.—Projected NEXT Thruster Life and Propellant Throughput versus NEXT Thruster 

Input Power, from NEXT Throttle Table. 
NASA/TM—2008-214986 14

100

1,000

10,000

100,000

1,000,000

1000 1500 2000 2500 3000 3500 4000 4500

Specific Impulse, seconds

Parameter

Life, khr

Throughput, kg Xe

Figure 13.—Projected NEXT Thruster Life and Propellant Throughput versus NEXT Thruster 

Specific Impulse, from NEXT Throttle Table. 

As such the reduced propellant throughput at these throttle points is of no consequence from a mission 

application stand point. Figures 12 and 13 plot the thruster life time and propellant throughput capability 

as a function of thruster input power and specific impulse respectively. As noted in from figure 12, the 

projected life time and propellant throughput of the thruster is >35 khr and >> 700 kg for most of the 

input power level, but experiences an inflection at <1.1 kW. From figure 13, this inflection point of 

reduced life time and throughput corresponds to specific impulse values less than about 2500 s. 

Summarizing the analyses performed to estimate NEXT thruster life capability versus throttle table 

throttle condition: 

9 The minimum life time is established by the ion optics accelerator electrode for all throttle conditions; 

9 The minimum life time is at full power, with the exception of operating points at thruster input powers 

≤665 W; 

9 For most of the throttle table the projected throughput capability is relatively constant at about 800 kg, 

corresponding to a total impulse of about 3×107

 N-s; 

9 The minimum throughput capability of the thruster is at low power; 

9 The cathode emitter wear mechanisms (tungstate formation and barium reactions, and barium 

depletion) and keeper erosion due to ion bombardment are not anticipated to be thruster life limiters. 

Table 4 lists demonstrated NEXT thruster performance over an extended-performance throttle table. 

These conditions—high power up to 14 kW, and high thrust-to-power - were investigated to understand 

the performance limits of the thruster (ref. 21); the NEXT thruster without modification is capable of 

operating over the conditions identified. Corresponding life times are projected in table 4 and these range 

from about 7.5 to 25 khr, with propellant throughputs from about 245 kg to > 800 kg. The total impulse 

and throughput capability of the thruster over most of the high-power portion of the throttle table are 

consistent with those values estimated for the nominal NEXT throttle table (table 3). Since the NEXT 

PPU is capable of operating at twice its maximum output beam current (7 A) at about half of its 

maximum output beam voltage (~1000 V), the NEXT PPU as-is can operate the thruster over all the high 
NASA/TM—2008-214986 15

thrust-to-power conditions identified in the extended-performance throttle table. It is also capable of 

operating at high power conditions with beam currents of 5.80 A and beam voltages ≤1179 V, and the 

7.0 A beam current condition at 1021 V beam voltage. Additional beam modules would need to be added 

to the PPU to capture the remaining extended-performance operating conditions. The PMS with 

appropriate re-sizing of the FCDs could operate over the entirety of the extended-performance throttle 

table. 

TABLE 4.—EXTENDED-PERFORMANCE THROTTLE TABLE: 

DEMONSTRATED THRUSTER PERFORMANCE (BOL) 

Vbps, 

V

F, 

mN 

Isp, 

s 

ηt Pin-t, kW MT, mg/s Ib, 

A 

Lt,

khr 

MT-TOT, 

kg Xe 

I, x107 Ns 

High Power Operation 

1800 389 4420 0.75 11.250 8.96 5.80 25.3 815 3.53 

1567 363 4130 0.74 9.930 8.96 5.80 25.3 815 3.30 

1396 343 3900 0.73 8.960 8.96 5.80 25.3 815 3.11 

1179 315 3585 0.72 7.730 8.96 5.80 25.3 815 2.86 

1021 293 3335 0.70 6.830 8.96 5.80 24.3 785 2.56 

1800 472 4460 0.76 13.645 10.79 7.04 18.7 730 3.18 

1567 441 4165 0.75 12.040 10.79 7.04 18.7 730 2.97 

1396 416 3930 0.74 10.865 10.79 7.04 18.7 730 2.80 

1179 382 3615 0.72 9.370 10.79 7.04 18.7 730 2.58 

1021 356 3360 0.71 8.285 10.79 7.04 9.66 375 1.24 

High Thrust-to-Power Operation 

700 147 2655 0.62 3.120 5.64 3.52 33.7 685 1.78 

700 167 2675 0.62 3.515 6.35 4.00 20.8 475 1.25 

700 187 2670 0.62 3.920 7.15 4.50 14.4 370 0.97 

700 208 2700 0.64 4.325 7.84 5.00 10.8 305 0.81 

700 228 2710 0.64 4.720 8.59 5.50 8.50 265 0.70 

700 240 2715 0.65 4.940 9.04 5.80 7.50 245 0.65 

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 Frontiersand Flagship-class studies. A summary of these findings are listed in table 5 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 5.—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 
NASA/TM—2008-214986 16

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; and identifying additional technology development and validation activities which 

may be of value in transitioning the TRL 6 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 6. 

TABLE 6.—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 semi-automatic 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 
NASA/TM—2008-214986 17

which may facilitate the transition of the NEXT IPS to flight and reduce barriers to 1st-user 

implementation. These tasks (see table 7) 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 flight-representative propulsion system test bed. It is anticipated that by taking these additional 

transition steps the 1st-user IPS implementation costs will be reduced.12 

TABLE 7.—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 2nd-generation 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 flight-system-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 

qualification-level environmental testing of key system elements are 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. 

 

12A 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—2008-214986 18

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: SingleString 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 MultiThruster 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. 

21. Patterson, M.J. “NEXT Study of Thruster Extended-Performance (NEXT STEP),” AIAA–2006–

4664, Joint Propulsion Conference, July 2006. 
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NEXT Ion Propulsion System Development Status and Performance 

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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 and prototype model system components, validating 

these through qualification-level and integrated system testing, and ensuring preparedness for transitioning to flight system 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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