John S. Snyder and John R. Anderson

Jet Propulsion Laboratory, Pasadena, California

Jonathan L. Van Noord and George C. Soulas

Glenn Research Center, Cleveland, Ohio

Environmental Testing of the NEXT PM1R

Ion Engine

NASA/TM—2008-215058

January 2008

IEPC–2007–276
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John S. Snyder and John R. Anderson

Jet Propulsion Laboratory, Pasadena, California

Jonathan L. Van Noord and George C. Soulas

Glenn Research Center, Cleveland, Ohio

Environmental Testing of the NEXT PM1R

Ion Engine

NASA/TM—2008-215058

January 2008

IEPC–2007–276

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Prepared for the

30th International Electric Propulsion Conference (IEPC 2007)

cosponsored by the Alta, S.p.A., Centrospazio, and the Electric Rocket Propulsion Society

Florence, Italy, September 17–20, 2007
Acknowledgments

The work described in this paper was carried out by the Jet Propulsion Laboratory, California Institute of Technology, under a 

contract with the National Aeronautics and Space Administration. The PM1R Environmental Test program, funded by NASA’s 

In-Space Propulsion Technology Program, was successfully accomplished through the efforts of a multiorganizational team. 

The work was performed under the NEXT Project, led by the NASA Glenn Research Center, with Scott Benson as the Project 

Manager and Mike Patterson as the Principal Investigator. Allison Owens and Ray Swindlehurst of the JPL Electric Propulsion 

Group were responsible for designing, confi guring, and operating the test facility. Rich Ebner, Tim Werner, and Doug Perry of 

the JPL Environmental Test Lab were responsible for setup, instrumentation, and performing the vibration test. Their efforts 

and contributions are greatly appreciated.

Available from

NASA Center for Aerospace Information

7115 Standard Drive

Hanover, MD 21076–1320

National Technical Information Service

5285 Port Royal Road

Springfi eld, VA 22161

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

Level of Review: This material has been technically reviewed by technical management. 

This report is a formal draft or working 

paper, intended to solicit comments and 

ideas from a technical peer group.
NASA/TM—2008-215058 1

Environmental Testing of the NEXT PM1R Ion Engine 

John Steven Snyder and John R. Anderson 

Jet Propulsion Laboratory 

Pasadena, California 91109 

Jonathan L. Van Noord and George C. Soulas 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Abstract 

The NEXT propulsion system is an advanced ion propulsion system presently under development that 

is oriented towards robotic exploration of the solar system using solar electric power. The subsystem 

includes an ion engine, power processing unit, feed system components, and thruster gimbal. The 

Prototype Model engine PM1 was subjected to qualification-level environmental testing in 2006 to 

demonstrate compatibility with environments representative of anticipated mission requirements. 

Although the testing was largely successful, several issues were identified including the fragmentation of 

potting cement on the discharge and neutralizer cathode heater terminations during vibration which led to 

abbreviated thermal testing, and generation of particulate contamination from manufacturing processes 

and engine materials. The engine was reworked to address most of these findings, renamed PM1R, and 

the environmental test sequence was repeated. Thruster functional testing was performed before and after 

the vibration and thermal-vacuum tests. Random vibration testing, conducted with the thruster mated to 

the breadboard gimbal, was executed at 10.0 Grms for 2 min in each of three axes. Thermal-vacuum testing 

included three thermal cycles from –120 to 215 °C with hot engine re-starts. Thruster performance was 

nominal throughout the test program, with minor variations in a few engine operating parameters likely 

caused by facility effects. There were no significant changes in engine performance as characterized by 

engine operating parameters, ion optics performance measurements, and beam current density 

measurements, indicating no significant changes to the hardware as a result of the environmental testing. 

The NEXT PM1R engine and the breadboard gimbal were found to be well-designed against 

environmental requirements based on the results reported herein. The redesigned cathode heater 

terminations successfully survived the vibration environments. Based on the results of this test program 

and confidence in the engineering solutions available for the remaining findings of the first test program, 

specifically the particulate contamination, the hardware environmental qualification program can proceed 

with confidence. 

I. Introduction 

NASA’s Evolutionary Xenon Thruster (NEXT) is an advanced ion propulsion system developed by a 

team led by the NASA Glenn Research Center (GRC) that is oriented towards robotic exploration of the 

solar system using solar electric power (ref. 1). Potential mission destinations that could benefit from a 

NEXT solar electric propulsion system include inner planets, small bodies, and outer planets and their 

moons (ref. 2). This range of robotic exploration missions generally calls for ion propulsion systems with 

deep throttling capability and system input power ranging from 5 to 25 kW, as referenced to solar array 

output at 1 Astronomical Unit (AU). In order to ensure that the NEXT system can successfully perform 

these missions, its components will be subjected to qualification-level testing that is representative of the 

anticipated environments. 
NASA/TM—2008-215058 2

The NEXT ion propulsion subsystem development includes a 7-kW high efficiency ion thruster, 

modular power processing unit (PPU), advanced propellant management system (PMS), lightweight 

gimbal, and software algorithms for a digital control interface unit (DCIU). Individual components in a 

subsystem such as this are typically subjected to separate environmental test sequences, partially due to 

the different environments each will experience on a spacecraft. For the test program discussed in this 

paper, the thruster and gimbal were tested together as an assembly in vibration environments. The thruster 

was tested alone in thermal environments. 

The NEXT ion thruster is a 7-kW, 36-cm-beam-diameter thruster with significant heritage to the 

2.3-kW, 30-cm NSTAR engine which flew on the Deep Space 1 mission (ref. 3) and is an integral part of 

the Dawn spacecraft scheduled for launch in late 2007 (ref. 4). Initial thruster development work was led 

by GRC and included significant performance testing (ref. 5), a 2000-hr wear test (ref. 6), an integrated 

test with other subsystem components (ref. 7), and more recently an ongoing wear test (ref. 8) and a 

multi-thruster array test (ref. 9). The thruster technology was transferred to NEXT industry partner 

Aerojet, which designed and fabricated a flight-like Prototype Model (PM) thruster (ref. 10). Following 

performance acceptance testing at GRC (ref. 11), the PM1 ion engine was sent to Jet Propulsion 

Laboratory (JPL) for extensive environmental testing. First the engine was subjected to a thermal balance 

test (ref. 12) in which the engine, instrumented with tens of thermocouples, was tested under a variety of 

operating conditions and thermal environments to characterize engine temperatures and thermal design 

margins. Data from that test were used in developing and validating an engine thermal model (ref. 13). 

Following the thermal balance test the PM1 engine was subjected to an environmental test program 

(ref. 14) in the latter half of 2006. Although the testing was largely successful there were several findings. 

The major issues encountered during the vibration test sequence were discharge and neutralizer heater 

impedance variations, particulate contamination generation, and fastener backouts on both the thruster 

and gimbal. Significant particulate contamination inside the engine was observed and was traced to 

fragmentation of the potting cement on the cathode heater terminations, nutplate wear, residue remaining 

inside the engine from manufacturing processes, and fraying fiberglass wire insulation. In many instances 

the amount of contamination generated was in excess of typical spacecraft cleanliness requirements. 

Additionally, during the third and final thermal cycle of the thermal-vacuum testing the neutralizer heater 

failed to pass current, most likely related to the impedance anomalies observed in vibration testing. 

Despite these findings, the thruster performance was nominal throughout the test program. There were 

noticeable changes in some thruster operational parameters, likely related to modest differences in tank 

pressure between different test phases, but no indication of a major change in thruster performance or 

operation resulting from exposure to the vibration and thermal environments. Similarly, there was no 

major change in ion optics performance or beam current density data. Although neutralizer cathode 

ignition times were consistent and showed little dependence on temperature, the discharge cathode 

ignition times showed a strong dependence on temperature. 

Most of the findings of the PM1 environmental test program were addressed during engine re-design 

activities performed at Aerojet, GRC, and JPL in late 2006 and early 2007 (ref. 15). The most significant 

work was focused on structural re-design of the cathode heater electrical terminations. Fastener loosening, 

electron backstreaming, nut plate galling, and fraying fiberglass wire insulation were also addressed. 

Complete resolution of all contamination issues was not possible with the reworked hardware, specifically 

the particulate contamination due to manufacturing processes, but engineering solutions are available and 

validation is planned with subsequent hardware builds. The re-worked engine, renamed PM1R, was 

successfully subjected to performance acceptance testing at GRC in mid-2007 and subsequently shipped 

to JPL for the repeated environmental testing which is reported here. 

The NEXT gimbal development was led by JPL. The breadboard gimbal is a lightweight, two-axis 

design using a four-bar linkage mechanism that provides a positive captured configuration for launch 

while providing for up to 20° of thrust angle articulation while in the operational configuration (ref. 16). 

JPL contracted with Swales Aerospace for the design and fabrication of two breadboard gimbals. The 

primarily-aluminum gimbal structure resembles an A-frame with titanium used for the flexures and the 

thruster mounts. An additional supporting bracket is required to support the mass of the gimbal motor. In 
NASA/TM—2008-215058 3

addition to the first thruster-gimbal vibration test, the breadboard gimbal was successfully subjected to 

vibration testing with a thruster mass model in an earlier test (ref. 17). 

The objectives of the environmental test program were as follows: 

• To validate the NEXT PM1R thruster and breadboard gimbal assembly to qualification-level 

dynamic environments, including: 

• Demonstrate thruster functionality pre- and post-vibe 

• Demonstrate gimbal functionality pre- and post-vibe 

• To validate the NEXT PM1R thruster to expected mission thermal environments, including 

standard qualification margins, including: 

• Expose the thruster to expected mission temperatures in cycles 

• Demonstrate thruster functionality pre- and post-test 

Environmental test requirements for the thruster and gimbal were determined from the NEXT project 

requirements, expected mission environments, and JPL assembly-level environmental verification 

policies and requirements. Discussion of the test requirements may be found in section II, along with test 

equipment and facility setup information, and test methods. Test results are presented in section III. 

II. Test Setup and Methods 

A. Test Article and Equipment 

1. Thruster and gimbal 

The thruster used for the environmental testing described herein was the re-worked NEXT Prototype 

Model 1 (PM1R) thruster. The thruster was originally manufactured and subsequently re-worked by 

Aerojet (Redmond, WA) under direction of the NASA Glenn Research Center. Significant changes in the 

PM thruster design compared to the previous generation of thruster designs include new material coatings 

to increase emissivity for enhanced thermal margin, more uniform ion optics apertures, and a graphite 

discharge cathode keeper to mitigate keeper erosion. Additional information about the PM1 engine, 

including performance data, is presented in reference 11. Brief descriptions of the modifications for the 

PM1R engine are discussed in reference 15. The cathode heater terminations were modified to be similar 

to the flight-proven Space Station Plasma Contactor design. Electron backstreaming was addressed 

through reducing the open-area fraction of the plasma screen. Fiberglass sleeving used for wire insulation 

was removed from the accessible internal wiring, but not at the main harness pass-through on the plasma 

screen. Loosened fasteners and nut plate galling were corrected with updated torque values and fastener 

types. 

Breadboard gimbal S/N 001 was used as the test article for the functional and vibration testing 

described herein. The breadboard gimbal was designed and fabricated to be flight-like, e.g., all fabricated 

parts were of JPL-approved materials with certifications. The motors chosen for the NEXT breadboard 

gimbal were ultra-high-vacuum class stepper motors. A flight unit would use space-rated versions with 

complete Quality Assurance documentation. The only functional difference between the two motor grades 

is the lower minimum temperature of the space-grade motor. The flight configuration of the gimbal has 

pinpullers specified as a launch lock for the actuator arm. This feature was not required for the breadboard 

gimbal, so a fixed-pin bracket was fabricated to replicate the launch lock function. All non-locking 

fasteners were locked or staked with Solithane, a flight-grade thread locking agent. All gimbal parts were 

demonstrated by analysis to meet allowable stress levels and minimum frequencies based on the predicted 

flight loads. Gimbal S/N 001 was vibrated under identical random vibration loads as to be used for the 

testing described herein with the PM1 thruster (ref. 14) and a thruster mass model (ref. 17). 
NASA/TM—2008-215058 4

2. Vacuum test facility 

The thruster test portions of the PM1R environmental test program were conducted in the JPL Patio 

Chamber facility. The vacuum chamber is 3 m in diameter and 8.6 m long, with ten cyropumps installed 

and operational for this testing. With the vacuum chamber configuration used for this test, specifically the 

inclusion of the large engine thermal shroud, the effective pumping speed was approximately 200,000 L/s 

on xenon. To minimize facility backsputter rates the interior of the vacuum facility is lined with graphite 

panels, which tended to absorb atmosphere gases and water vapor during the vacuum cycles of the testing. 

Modest tank pressure variations were observed as the panels outgassed during thruster operation. To 

mitigate this effect, a full-power thrusting operating phase was added prior to functional testing to bake 

out the panels. 

The thruster was installed in a 1.2-m diameter by 1.0-m long cylindrical thermal shroud located at one 

end of the vacuum facility. The downstream end of the thermal shroud was located 6.2 m from the 

downstream end of the vacuum facility. The thruster was mounted inside the shroud with the neutralizer 

keeper orifice plate located 5.7 cm from the downstream end of the shroud. The cylindrical and rear 

thermal shroud walls were equipped with liquid nitrogen tubing to provide environmental cooling of the 

test article. The opening at the front of the shroud was equipped with a door made of multi-layer 

insulation. This door was opened during thruster operation and could be closed to minimize thermal 

interaction with the vacuum facility during non-operational cold soak periods. 

The thermal shroud was also equipped with eight quartz infrared heat lamps to provide environmental 

heating to the thruster. The lamps were installed parallel to the thruster axis and spaced 45° apart 

azimuthally with one end of the lamp even with the front mask of PM1R. A temperature controller and 

SCR power controller were used for tight process control during temperature ramp and hold. A 

photograph of the test setup with the heat lamps operating is shown in figure 1. In contrast to the Thermal 

Development Test (ref. 12) where the heat lamps were controlled through the heat flux coupons mounted 

near the engine, in the thermal vacuum test the heat lamps were directly controlled through the 

temperature of gimbal pad A. Gimbal pad A was located at the 6 o’clock position as viewed in figure 1 

(i.e., opposite the neutralizer); pads B and C were located at the 10 and 2 o’clock positions, respectively. 

Gimbal pad A was chosen as the reference location because it was the coolest of the three gimbal pads 

during preliminary testing. 

Figure 1.—Test setup for functional and thermal vacuum testing. 
NASA/TM—2008-215058 5

The thermal shroud and facility were equipped with more than twenty thermocouples for process 

control and monitoring. The thruster propellant lines internal to the thermal shroud were wrapped with 

heater tape, instrumented with thermocouples, and wrapped in an MLI blanket to control the temperature 

of the lines during thruster cold soaks and periods prior to engine startup. 

Beam current density profiles were measured with two Faraday probes. The probes were installed on 

a stage that allowed them to translate through the thruster plume at axial distances between 0.045 and 

0.55 m downstream of the thruster. 

The environmental testing was performed using laboratory power supplies controlled and monitored 

by a data acquisition and control system. Dedicated cathode ignitor circuits were used for cathode ignition 

and grid recycling was handled by a custom recycle circuit. Flow and electrical instrumentation were 

calibrated prior to the testing. The data acquisition system recorded thruster currents, voltages, flow rates 

and temperatures, and facility pressure and temperatures at a user-specified rate. Data were typically 

recorded once a minute; however, during thruster starts or when thruster parameters were being varied the 

rate was often changed to once every ten seconds. The software used to record data was also used to 

control thruster power supplies and flow rates. Closed-loop control of flow rates and beam current was 

used, the latter by adjusting discharge current. 

3. Vibration test setup and facility 

The thruster was mated to the breadboard gimbal for vibration testing, as shown in the test setup 

photograph of figure 2. The thruster-gimbal assembly was mounted on a vibration fixture, comprised of a 

vibration test plate and three vibration test mounts, which provided the mechanical interface between the 

vibration table and the test article. The engine was fixed to the gimbal at each of the three gimbal legs. 

Each gimbal leg rested on a vibration test mount, and four piezoelectric three-axis force transducers were 

mounted between each test mount and the vibration test plate. These were used for vibration force 

limiting, discussed later. Gimbal legs were annotated the same as the thruster pads, i.e., leg A was located 

opposite the neutralizer. 

The test assembly was instrumented with a total of seventeen tri-axial and one uni-axial response 

accelerometers. All accelerometers were bonded to Kapton tape applied to the test article surfaces. Nine 

tri-axials were mounted on the vibration test mounts near the gimbal strut attachment points. Three triaxials were located on the side of the gimbal brackets that mount to the thruster, and three other tri-axials 

Figure 2.—Test setup for Z-axis vibration. 
NASA/TM—2008-215058 6

were mounted on the thruster directly below the interface with the gimbal bracket. The remaining two triaxials were mounted on the neutralizer and the neutralizer support pad, and the lone uni-axial 

accelerometer was mounted on the unperforated region of the accelerator grid directly opposite the 

neutralizer. 

Vibration testing was performed in the JPL Dynamics Environmental Test Facility. Vertical (i.e., Zaxis, or thrust axis) excitation was performed on a Kimball head expander on the large spacecraft 

vibration test stand. Lateral (X- and Y-axis) excitation was performed on a separate oil-lubricated slide 

plate. Application of vibration to the test article was controlled using one of JPL’s m+p vibration control 

systems. Vibration test control instrumentation consisted of two control accelerometers and one monitor 

accelerometer mounted to the vibration test plate. This was in conjunction with the signal from the twelve 

force transducers at the unit interface, controlling all random vibration tests using the extremal control 

mode. Data acquisition from control and response accelerometers was provided at a 20 kHz sampling 

rate. 

B. Test Methods 

The environmental test program was performed in the sequence of steps shown in table 1. 

Additionally, hardware inspection was performed prior to the first and after each of the test segments. 

This included visual inspection, electrical diagnostics, and in some cases disassembly of test hardware. 

Notable inspection results are described in the appropriate sections of this paper. Thruster and 

thruster/gimbal functional testing was performed prior to and after each test segment to demonstrate 

operation over requirements after exposure to environments. 

TABLE 1.—ENVIRONMENTAL TEST SEQUENCE 

Sequence Operation 

1 Pre-vibe thruster functional test 

2 Pre-vibe gimbal functional test 

3 Random vibration tests 

4 Post-vibe gimbal functional test 

5 Post-vibe thruster functional test 

6 Thermal vacuum test 

7 Post-thermal thruster functional test 

1. Thruster functional test 

Thruster functional testing was performed prior to the vibration testing to demonstrate nominal engine 

operation, including: engine performance, neutralizer characterization, electron backstreaming and 

perveance margins, double-to-single ion current ratios, and beam current density profile. Thruster 

functional testing was repeated after vibration testing and thermal-vacuum testing. 

Testing of the NEXT PM1R thruster was conducted using laboratory power supplies and a laboratory 

flow system. A total of four Functional Test Points (FTP) were used during the environmental testing; the 

thruster controlled parameters are listed in table 2 for those conditions as well as the startup conditions 

(i.e., Discharge-only). Seven parameters were specified for a given operating condition: the propellant 

flow rates to the discharge chamber, cathode, and neutralizer; the extraction parameters beam current, 

beam voltage, and accelerator grid voltage; and the neutralizer keeper current. Note that in practice it was 

the screen power supply voltage that was controlled and the beam voltage was calculated from the sum of 

the power supply voltage and the neutralizer floating voltage. Beam current was controlled by adjusting 

the discharge current, often performed automatically with closed-loop control through the data acquisition 

and control system. Operating conditions are referenced in this report by the combination of beam current 

and voltage. 
NASA/TM—2008-215058 7

TABLE 2.—THRUSTER CONTROLLED PARAMETERS FOR FUNCTIONAL TESTING 

Control parameter Discharge-only FTP 1 FTP 2 FTP 3 FTP 4 

Beam current, (A) ---- 1.20 2.00 3.52 3.52 

Beam voltage, (V) ---- 679 1179 1179 1800 

Accelerator voltage, (V) ---- –115 –200 –200 –210 

Neutralizer keeper current, (A) 3.0 3.0 3.0 3.0 3.0 

Main flow rate, (sccm) 14.23 14.23 25.79 49.64 49.64 

Cathode flow rate, (sccm) 3.57 3.57 3.87 4.87 4.87 

Neutralizer flow rate, (sccm) 6.00 3.00 2.50 4.01 4.01 

Thruster operating parameters were continuously recorded by the data acquisition system during the 

thruster functional performance testing. In addition to the controlled parameters, the following parameters 

were recorded: accelerator grid current, discharge current and voltage, cathode keeper voltage, neutralizer 

keeper voltage, neutralizer-common-to-ground voltage, and tank pressure. Common calculated or derived 

parameters were also determined: thrust, specific impulse, total efficiency, and beam ion energy cost (also 

called discharge losses). Floating (i.e., neutralizer coupling) voltages were in the range of –9.9 to –11.7 V 

for this testing. 

Ion optics performance was measured using standard procedures. Perveance measurements were 

made by holding the beam current and accelerator grid voltage constant while varying the screen grid 

voltage and recording the accelerator grid current. The discharge current was adjusted in this case to 

maintain constant beam current. The perveance limit was defined as the total voltage at which the rate-ofchange of accelerator grid current was –0.02 mA/V. Electron backstreaming (EBS) onset was determined 

by reducing the magnitude of the accelerator grid voltage at constant discharge current and monitoring the 

beam current. A 1-mA increase in the beam current from the lowest current achieved during the 

measurement defined the EBS limit. In practice, the voltage drop across the calibrated beam current shunt 

resistor was used to determine this limit. 

Neutralizer characterization was performed in the discharge-only mode conditions shown in table 2 

by decreasing the neutralizer flow rate until the neutralizer transitioned from spot to plume mode. Plume 

mode, characterized by electrical oscillations in the neutralizer keeper circuit, was defined as reaching or 

exceeding 5 V peak-to-peak oscillations measured between neutralizer keeper and neutralizer common. 

2. Gimbal functional test 

Gimbal functional testing was performed with the thruster integrated onto the gimbal and the 

assembly mounted to the vibration test fixturing. Since the gimbal is a zero-G design that cannot 

adequately support and drive the thruster mass in a one-G environment, the thruster-gimbal assembly was 

fitted with a system of weights and pulleys for functional testing to offload the thruster mass. Offloading 

was applied equally at each gimbal pad through a fastener attached to each gimbal thruster bracket as 

shown in figure 3. The gimbal was driven via a custom software program and dedicated motor control 

hardware. 

Functional testing was accomplished by driving the gimbal to a range of nineteen positions which 

described and circumscribed the range of motion. At several of the sequence steps, an inclinometer was 

used to measure the gimbal pitch in two orthogonal axes. Accuracy of the inclinometer measurement was 

estimated to be ±0.5° and all measurements were rounded to the nearest degree. After returning to the 

stowed position, the pin-puller mockups were re-inserted into the assembly and their fit used to determine 

if the stepper motors returned to the correct final position (i.e., there were no missed steps during the 

functional test). 
NASA/TM—2008-215058 8

Figure 3.—Gimbal functional test. 

3. Vibration test 

The random vibration specification for the NEXT thruster-gimbal assembly was based on Delta-class 

launch vehicles and typical locations of electric propulsion hardware on spacecraft. The qualificationlevel specification, shown in table 3, has a total level of 10.0 Grms and calls for random vibration testing in 

each of three orthogonal axes for a duration of 2 min per axis. Sine surveys were performed before and 

after each full-level random vibration test to identify hardware changes caused by testing. The surveys 

were run at a load of 0.25 G0-pk from 5 to 2000 Hz at a sweep rate of two octaves per minute. 

TABLE 3.—QUALIFICATION-LEVEL RANDOM VIBRATION REQUIREMENTS 

Frequency, 

(Hz) 

Level 

20 0.04 g2/Hz 

20 to 50 +3 dB/octave 

50 to 600 0.1 g2/Hz 

600 to 2,000 –6 dB/Octave 

2000 0.009 g2/Hz 

Overall 10.0 Grms

Duration: 2 min./axis, 3 orthogonal axes 

Vibration testing of the NEXT thruster-gimbal assembly also incorporated the practice of forcelimited vibration testing which is used for most vibration tests performed at JPL. Force-limiting during 

vibration testing guards against artificial test failures caused by overtesting, a result of the infinite 

mechanical impedance of the shaker and the use of only acceleration-based control (ref. 18). In this 

situation, the reaction forces at the fixture/engine interface can become unrealistically high compared to a 

flight environment at test article resonances. In a force-limited test, real-time force measurement and 

limiting is performed to notch the input acceleration spectrum. 

Vibration testing in the Z-axis (i.e., thrust axis) was performed first, followed by the lateral axes Y 

(i.e., axis intersecting both the discharge chamber and neutralizer centerlines) and finally X. After the 

Z-axis testing was completed the test assembly including vibration fixturing was removed from the head 
NASA/TM—2008-215058 9

expander and was attached to the slip plate for lateral axis testing. For all axes, the first activity was a sine 

survey of the assembly. This was followed by a random vibration test at a level of –18 dB with respect to 

the full random vibration loads shown in table 3 in order to verify instrumentation operation and tune the 

vibration force-limiting algorithms. The full-level random vibration test was then performed, followed by 

a final sine survey. Physical inspections of the test assembly were performed after each of these tests. 

4. Thermal vacuum test 

Requirements for the thermal vacuum test were derived from mission studies and simple spacecraft 

configuration models. The NEXT system design has been driven by Deep Space Design Reference 

Missions as well as by Discovery and New Frontiers class missions (ref. 19). Preliminary modeling of the 

NEXT PM-design thruster on a sample spacecraft configuration showed that the most aggressive thermal 

conditions were projected to occur during Venus gravity assists on trajectories to Saturn and Neptune, 

specifically at 0.85 A.U. and with a solar incidence angle of 38° from the thruster exit plane (ref. 10). A 

more rigorous thermal model of the NEXT thruster was recently developed in conjunction with the PM1 

Thermal Development Test (ref. 12) and validated with data from that test (ref. 13). 

A departure from standard practice was made with the selection of reference temperature location on 

the NEXT engine. On the Deep Space 1 mission, a location on the front mask of the NSTAR ion engine 

was used as the reference temperature location for environmental testing and flight telemetry (ref. 20). It 

was very difficult, however, to correlate the reference location temperature with the internal thruster 

temperatures because the mask was not well coupled to the engine through thermal conduction and 

knowledge of the surface optical properties was poor. Evaluation of data from the PM1 Thermal 

Development Test combined with the thermal model of the engine suggested that the engine gimbal pads 

would be a more appropriate location for temperature reference. 

The NEXT PM1R thermal model, updated to reflect pertinent design changes, was employed in a 

simple spacecraft configuration model at the projected mission worst-case solar irradiance to determine 

maximum expected gimbal pad temperatures in this environment. With the reduced open-area fraction of 

the plasma screen mesh compared to PM1, the maximum expected temperature increased from 187 °C 

(see ref. 13) to 195 °C. The thermal-vacuum test maximum temperature for PM1R was thus 215 °C 

including the standard flight temperature margin. The cold temperature limit for NEXT thermal-vacuum 

testing was not derived from corresponding mission analyses and thermal modeling; ion thrusters can 

survive much lower temperatures than the –100 °C levels typically required for flight systems. For the 

testing described herein, a cold temperature of –120 °C was chosen which was 11 °C colder than the Deep 

Space 1 qualification requirement for the NSTAR ion engine (ref. 20). 

The plan for the thermal vacuum test called for three identical thermal cycles. The cycles consisted of: 

1. 2-hr minimum cold soak after reaching the cold temperature limit of –120 °C. 

2. Short-duration operation in discharge-only mode at full-power conditions. 

3. Full-power operation ramp to the hot temperature limit of 215 °C. 

4. 4 hr of operation at full power (3.52 A and 1800 V). 

5. 4 hr of operation at mid power (3.52 A and 1179 V). 

6. Turn off engine, then immediately begin cathode ignition sequence. 

7. Short-duration operation in discharge-only mode, then at full power conditions. 

8. Turn off engine and cool for next cycle. 

A schematic of the test profile is shown in figure 4. In this test profile, three cold starts and three hot 

starts were included. The planned total cumulative cold soak time was 6 hr and total cumulative hot 

operation time was 24.5 hr. All discharge-only operation during thermal-vacuum testing was performed at 

discharge conditions equivalent to full power operation (i.e., 3.52 A and 1800 V). Engine operating 

parameters were recorded continually throughout the test as described in section II. After approximately 

3 hr of operation at each of the two beam extraction operating points, electron backstreaming and beam 

current density data were acquired. 
NASA/TM—2008-215058 10

200

100

0

-100

-200

Temperature (oC)

0 10 20 30 40 50 60 70 80 90 100

Time (hours)

 Thruster ON

 Thruster OFF

Times are approximate

Figure 4.—Thermal vacuum test plan temperature profile. 

III. Test Results 

A. Pre-Vibration Thruster Functional Test 

Thruster functional testing was performed at the beginning of the environmental test program to 

establish a baseline against which post-environmental performance could be compared. Non-controlled 

and derived engine performance parameters for the highest power and a mid power operating condition 

are compared in table 4 for the PM1 (ref. 14) and reworked PM1R pre-vibration functional tests. Note, as 

will be seen in the data presented in a subsequent section of this paper, the differences between the PM1 

and PM1R data are not significantly different from the scatter typically seen between different tests with 

the same engine. The modest differences in the discharge parameters and accelerator grid current are 

likely due to differences in the tank pressures, which were higher by about 10 percent in the PM1 testing. 

Note that at higher background pressures there is more mass ingestion into the discharge chamber, and 

standard operating procedure for the NEXT test program is not to correct the main flow into the thruster 

for ingestion, which can affect the calculated beam ion energy cost and total thruster power through the 

measured discharge power. Apart from these noted parameters, there was no significant difference 

observed in the engine performance data. Ion optics performance margins were also comparable, within 

the typical measurement error and data scatter. As expected, the rework done on the thruster had no 

apparent effect on the measured engine performance. 

TABLE 4.—COMPARISON OF PRE-VIBRATION ENGINE PERFORMANCE DATA 

FOR PM1 (REF. 14) AND PM1R THRUSTERS

 3.52 A, 1800 V 2.0 A, 1179 V 

 PM1R PM1 PM1R PM1 

Discharge current, (A) 18.6 18.1 14.2 13.9 

Discharge voltage, (V) 24.4 23.2 24.2 24.5 

Cathode keeper voltage, (V) 5.5 5.4 3.6 3.9 

Accelerator grid current, (mA) 25.9 20.8 10.8 7.2 

Neutralizer keeper voltage, (V) 11.8 12.0 13.5 13.3 

Power, (W) 6890 6860 2750 2760 

Thrust, (mN) 237 238 108 109 

Specific impulse, (sec) 4180 4190 3470 3500 

Total efficiency 0.705 0.712 0.668 0.677 

Perveance margin, (V) 1150 1156 708 686 

Electron backstreaming margin, (V) 43 49 101 101 
NASA/TM—2008-215058 11

B. Pre-Vibration Gimbal Functional Test 

Starting from the stowed position, the gimbal was first driven to the “clear latch” position to ensure 

that the thruster pins cleared the gimbal A-frames acceptably. The gimbal was then driven to each of the 

positions shown in table 5 and the inclination was measured at gimbal thruster bracket A (i.e., the gimbal 

bracket attached to the thruster, located opposite the neutralizer) in the radial and tangential directions. 

The total range of the gimbal at this bracket was measured as –19°/20° in the radial direction and 

16°/–17° in the tangential direction. The full nineteen-step drive program was then successfully 

performed to exercise the gimbal motion with the attached thruster. Following completion of the 

sequence, the pin-puller mockups were re-inserted into the gimbal demonstrating successful and complete 

return to the stowed and latched position. 

TABLE 5.—PRE-VIBE GIMBAL ARTICULATION MEASUREMENT RESULTS

Position name Gimbal pad A 

elevation, 

(cm) 

Radial 

inclination, 

° 

Tangential 

inclination, 

° 

Stowed, Latched 0 0 0 

Clear latch 2 0 0 

Mid plane 10 0 0 

Max A, Level B, Level C 17 –8 0 

Max A, Min B,C 17 –19 0 

Min A, Max B,C 1 20 0 

Mid plane 10 0 0 

Level A, Min B, Max C 10 0 16 

Level A, Max B, Min C 10 –0 –17 

Stowed, Latched 0 0 0 

C. Thruster-Gimbal Assembly Vibration Test 

1. Z-axis testing 

Dynamic testing began with the Z-axis (i.e., thrust-axis) testing, shown in the photograph of figure 2. 

Initial tests and then the 2-min full random vibration test were completed without incident. The vibration 

test inputs, which were notched automatically and in real time during the test according to the force limit 

specifications, are shown in figure 5. Significant notches were observed at 100, 760, and 1130 Hz, similar 

to the behavior of the previous test (ref. 14). 

Comparison of the pre- and post-vibe sine surveys showed significant changes on all in-axis 

accelerometers on the thruster and gimbal near the fundamental mode. A typical response is shown in 

figure 6 for the accelerometer mounted on thruster gimbal pad B, just underneath the gimbal bracket. A 

frequency shift of roughly 10 percent was observed near 100 Hz with little-to-no evidence of a change in 

response damping, for all accelerometers on the thruster and gimbal and for the measured in-axis force 

sum. The frequency shift is larger than what is nominally observed in run-to-run comparisons and merits 

additional investigation. 
NASA/TM—2008-215058 12

10-3

10-2

10-1

100

Power Spectral Density (g2/Hz)

10 100 1000

Frequency (Hz)

 Specification (10.0 Grms)

 Control (9.2 Grms)

Figure 5.—Comparison of random vibration specification with force-limited 

test inputs, Z-axis excitation. 

10-3

10-2

10-1

100

101

Accelerometer Response (G's)

10 100 1000

Frequency (Hz)

Accelerometer A2Z

 Pre-Vibe

 Post-Vibe

Figure 6.—Comparison of sine survey results before and after full-level Z-axis 

random vibration, thruster gimbal pad B. 

In contrast to the changes observed on the thruster and gimbal, there was little change observed in the 

data from accelerometers located on the vibration test mounts, as shown in the representative data of 

figure 7. These results collectively indicate that changes occurred in the test article and not at the 

structural bolting interface between the vibration fixturing and the vibe table. 

Inspection of the test assembly immediately after the Z-axis testing did not reveal any clear causes for 

the sine survey shifts observed in the data. Frequency shifting and some damping have been observed in 

previous Z-axis testing with this gimbal (refs. 14 and 17) and it is believed that these shifts are due to 

settling in the hardware at the capture interface between the thruster pin and the gimbal frame but this has 

not yet been proven. At present, additional analyses are underway to understand this issue and determine 

if any design changes are necessary. Nevertheless, as will be shown, there was no observed effect on the 

ability of the gimbal or thruster to pass post-vibe functional tests. 
NASA/TM—2008-215058 13

10-2

10-1

100

101

Accelerometer Response (G's)

10 100 1000

Frequency (Hz)

Accelerometer A16Z

 Pre-vibe

 Post-vibe

Figure 7.—Comparison of sine survey results before and after full-level 

Z-axis random vibration, vibe test mount C, near free leg. 

Figure 8.—Test setup for X-axis vibration. 

2. Lateral axis testing 

Dynamic testing continued with the Y-axis testing, which was completed without incident. Following 

the Y-axis testing the test assembly was rotated 90° to perform the X-axis random vibration testing, 

shown in the test setup of figure 8. The X-axis test sequence including the 2-min full random vibration 

test was also completed without incident. The vibration test inputs for both X and Y excitation are shown 

in figure 9. Significant notches in the control acceleration spectrum were observed at 76 and 400 Hz for 

the Y-axis excitation, and at 80 and 380 Hz for X-axis excitation. 

Comparison of the Y-axis pre- and post-vibe sine surveys showed only minor changes in the test 

assembly near the fundamental mode in the frequency range of 50 to 100 Hz. Larger differences were 

seen in the X-axis comparisons; typical data are shown in figure 10, similar differences were observed 

throughout the other accelerometers on the test article. A sizable narrow-band peak near the 
NASA/TM—2008-215058 14

10-3

10-2

10-1

100

Power Spectral Density (g2/Hz)

10 100 1000

Frequency (Hz)

 Specification (10.0 Grms)

 X-Axis Control (9.8 Grms)

 Y-Axis Control (9.6 Grms)

Figure 9.—Comparison of random vibration specification with force-limited 

test inputs, lateral axis excitation. 

10-3

10-2

10-1

100

101

Accelerometer Response (G's)

10 100 1000

Frequency (Hz)

Accelerometer A1X

 Pre-Vibe

 Post-Vibe

Figure 10.—Comparison of sine survey results before and after full-level 

X-axis random vibration, thruster gimbal pad A. 

fundamental mode that was observed before the full random vibration test did not appear afterward. This 

peak was also observed in the PM1 vibration test of 2006, but did not decrease as seen in figure 10. As for 

the Z-axis results, the accelerometers on the vibration test mounts did not show any significant difference 

for either Y-axis or X-axis testing. This result is an indication that changes occurred in the test article and 

not at the structural bolting interface between the vibration fixturing and the vibe table. As noted earlier, 

these sine survey changes are presently under investigation. 
NASA/TM—2008-215058 15

3. Post-vibration inspection results 

The thruster/gimbal assembly was given an extensive inspection following the completion of the 

vibration testing. The interior of the thruster discharge chamber was visually inspected through the ion 

optics and a small amount of particulate contamination was observed. This was not unexpected, as a 

significant amount of residue from a mechanical fabrication treatment was observed following the 

previous engine vibration test (ref. 14), and was difficult to remove entirely during the re-work process. 

Although thruster pin rotation had been observed in previous testing, most notably in the mass model 

vibration test, with the additional staking compound applied to the pin in the PM1R testing there was no 

observable pin rotation in this vibration test program. 

Inspection of the bottom of the plasma screen cone revealed a single, relatively large piece of debris. 

The debris was retrieved from the engine by removing a single section of the plasma screen and appeared 

to be fibrous material. Analysis of the debris positively identified the material as a match to the fiberglass 

wire sleeving originally used in the PM1 engine. A small amount of the sleeving material near the harness 

passthrough on the plasma screen was not removed during the engine re-work and was identified as the 

likely source. 

D. Post-Vibration Gimbal Functional Test 

Gimbal functional testing was repeated in the same manner that it had been performed in the previbration characterization. The gimbal was driven to each of the positions shown in table 5 and the 

inclination was measured at gimbal thruster bracket A in the radial and tangential directions. All 

measurements were identical to the pre-vibe measurements of table 5 with the exception of the tangential 

inclination in the “Level A, Min B, Max C” step which was 16° in the pre-vibe test and 17° in the postvibe test. This is within the uncertainty of the measurement and the rounding applied to the data. The total 

range of the gimbal at this bracket was measured as –19°/20° in the radial direction and ±17° in the 

tangential direction, identical to the pre-vibration measurement within uncertainty. The full nineteen-step 

drive program was then successfully performed to exercise the gimbal motion with the attached thruster. 

Following completion of the sequence, the pin-puller mockups were re-inserted into the gimbal 

demonstrating successful and complete return to the stowed and latched position. There was no indication 

from inspection or test that the gimbal performed any differently after vibration than it did before 

vibration. 

E. Post-Vibration Thruster Functional Test 

The post-vibration thruster functional testing was conducted in the same manner as the pre-vibration 

testing. There were no significant differences in engine performance as compared to the pre-vibration 

testing. Detailed comparisons of performance data are shown in section G. 

F. Thruster Thermal-Vacuum Test 

The first thermal cycle was performed following engine cool down initiated the previous day. The 

cathode ignition process was initiated approximately seven hours after the reference temperature cooled 

to the target of –120 °C. The engine started nominally and was brought to full power per the test plan. 

Heat lamp power was then turned on and the temperature reference location ramped to the 215 °C control 

temperature in 69 min. Thruster operation was held steady at full power (3.52 A, 1800 V) and at the 

control temperature for 241 min before transitioning to the lower power point (3.52 A, 1179 V). After 

240 min of operation at this condition the thruster was shut down and the cathode heaters immediately 

turned on for the hot start. The hot start was nominal and the engine operated in discharge-only mode per 

the test plan and then at full power before finally turning the engine off. The engine and thermal shroud 
NASA/TM—2008-215058 16

250

200

150

100

50

0

-50

-100

-150

Temperature (oC)

8/13/2007 8/14/2007 8/15/2007 8/16/2007

Runtime

8/20/2007 8/21/2007

Gimbal Pad C

Figure 11.—Thermal vacuum testing temperature profile. 

were allowed to cool for the next thermal cycle. Performance data obtained during the thermal cycles are 

discussed in a subsequent section. 

The second and third thermal cycles were also performed nominally per the test plan. The cathode 

ignition process was initiated approximately four hours after the reference temperature cooled to the 

target of –120 °C for the second cycle (4.5 hr for the third cycle). The engine started nominally and was 

brought to full power per the test plan. Heat lamp power was turned on and the temperature reference 

location ramped to the 215 °C control temperature in 69 min (70 min for the third cycle). Thruster 

operation was held steady at full power and the control temperature for 240 min (for both cycles) before 

transitioning to the lower power. After 242 min of operation at this condition (240 min for the third cycle) 

the thruster was shut down and the cathode heaters immediately turned on for the hot start. The hot start 

was nominal and the engine operated in discharge-only mode per the test plan then at full power before 

finally turning the engine off (for both cycles). There were no issues with thruster operation during any of 

the thermal-vacuum testing. 

The temperature history for all of the thermal-vacuum testing is shown in figure 11. All three gimbal 

pad temperatures were within a few degrees of each other during the testing, hence only gimbal pad C is 

shown in the figure for clarity. Note that the coolest of the three gimbal pads was chosen for the control 

thermocouple. In the hot equilibrium condition, gimbal pad C measured 218 °C. 

G. Thruster Functional Testing 

Thruster functional testing was performed in the same manner for all functional tests during the 

environmental test program. Non-controlled and derived engine performance parameters for the lowerpower operating conditions are compared for all functional tests in table 6. It can be seen that there were 

no major differences in performance throughout the environmental test program at these power levels. 

Minor differences of up to a few percent were observed in the non-controlled operating parameters, with 

discharge current variations of several percent at the lowest power. Discharge chamber parameters and 

accelerator grid current were likely influenced by the tank pressure differences of up to 10 percent 

between tests (20 percent for the first set of tests at the lowest power) caused by graphite panel gas 

absorption/desorption during the vacuum cycles and rapid tests. Changes in discharge chamber neutral 

density caused by ingested flow affect discharge performance through the primary electron loss rate and 

electron temperature, for example (ref. 21). Neutral density near the accelerator grid is also a significant 

factor in charge-exchange ion production and accelerator grid current. Slight variations in the derived 

parameters, such as the beam ion energy cost, reflect the differences in measured parameters. The largest 

effect on the efficiency at the lowest power was the change in discharge power. 
NASA/TM—2008-215058 17

Engine performance for the 3.52 A, 1179 V operating condition is shown in table 7, which includes 

performance from the three thermal cycles where the engine was operating at elevated temperatures. 

Again, no major differences in performance were observed with the exception of the accelerator grid 

current which is almost certainly due to tank pressure variations (it will be shown later that optics 

performance was essentially unchanged during the testing). Performance data for the full-power operating 

condition are summarized in table 8. As for the mid-power cases shown in table 7, performance was 

largely the same with the exception of variations in the accelerator grid current due to tank pressure. Note 

that outgassing from the thruster and/or thermal shroud may have also contributed to the generally-higher 

measured tank pressures during the thermal cycle testing at elevated temperatures. 

TABLE 6.—ENGINE PERFORMANCE FOR THE LOWER-POWER FUNCTIONAL TEST POINTS

1.2 A, 679 V 2.0 A, 1179 V 

Pre-vibe Post-vibe Postthermal Pre-vibe Post-vibe Postthermal 

Discharge current, (A) 10.4 10.9 10.1 14.2 14.4 14.3 

Discharge voltage, (V) 26.1 26.3 26.2 24.2 24.4 24.1 

Cathode keeper voltage, (V) 3.4 3.1 2.9 3.6 3.4 3.0 

Accelerator grid current, (mA) 5.9 4.1 3.9 10.8 8.2 7.5 

Neutralizer keeper voltage, (V) 13.5 13.9 13.8 13.5 13.5 13.4 

Power, (W) 1130 1140 1120 2750 2760 2750 

Tank pressure, (Torr Xe) 2.0E-06 1.5E-06 1.5E-06 2.7E-06 2.4E-06 2.3E-06 

Thrust, (mN) 49.2 49.2 49.1 108 108 108 

Specific impulse, (sec) 2440 2440 2440 3470 3470 3480 

Total efficiency 0.521 0.515 0.524 0.668 0.669 0.672 

Beam ion energy cost, (eV/ion) 227 238 221 173 176 172 

TABLE 7.—ENGINE PERFORMANCE FOR THE 3.52 A, 1179 V OPERATING CONDITION 

Pre-vibe Post-vibe Thermal 

cycle #1 

Thermal 

cycle #2 

Thermal 

cycle #3 

Postthermal 

Discharge current, (A) 20.1 20.1 19.7 19.8 19.9 20.2 

Discharge voltage, (V) 25.2 25.4 24.9 24.9 25.0 25.0 

Cathode keeper voltage, (V) 5.5 5.5 5.4 5.4 5.4 5.1 

Accelerator grid current, (mA) 24.6 20.4 27.8 24.5 22.4 19.3 

Neutralizer keeper voltage, (V) 11.7 11.7 12.0 11.8 12.1 11.8 

Power, (W) 4710 4710 4700 4700 4700 4700 

Tank pressure, (Torr Xe) 4.8E-06 4.2E-06 4.7E-06 4.4E-06 4.4E-06 4.2E-06 

Thrust, (mN) 191 191 191 191 191 191 

Specific impulse, (sec) 3380 3370 3380 3380 3380 3370 

Total efficiency 0.672 0.672 0.675 0.675 0.675 0.672 

Beam ion energy cost, (eV/ion) 144 145 139 140 141 144 

TABLE 8.—ENGINE PERFORMANCE FOR THE 3.52 A, 1800 V OPERATING CONDITION 

Pre-vibe Post-vibe Thermal 

cycle #1 

Thermal 

cycle #2 

Thermal 

cycle #3 

Postthermal 

Discharge current, (A) 18.6 18.6 18.5 18.7 18.8 18.9 

Discharge voltage, (V) 24.4 24.4 24.3 24.3 24.4 24.6 

Cathode keeper voltage, (V) 5.5 5.4 5.3 5.2 5.3 5.1 

Accelerator grid current, (mA) 25.9 22.2 29.4 24.1 21.2 18.7 

Neutralizer keeper voltage, (V) 11.8 11.8 12.1 11.8 11.8 11.8 

Power, (W) 6890 6870 6880 6870 6870 6880 

Tank pressure, (Torr Xe) 5.1E-06 4.8E-06 5.2E-06 4.6E-06 4.4E-06 4.3E-06 

Thrust, (mN) 237 237 237 237 237 237 

Specific impulse, (sec) 4180 4180 4180 4180 4180 4180 

Total efficiency 0.705 0.705 0.705 0.706 0.706 0.706 

Beam ion energy cost, (eV/ion) 129 129 128 129 131 133 
NASA/TM—2008-215058 18

Ion optics performance measurements are compared for all functional testing in table 9. Note that 

during the thermal vacuum testing perveance measurements were not made, nor were electron 

backstreaming (EBS) measurements made at the two lower power conditions. There were no measurable 

changes observed in optics performance as a result of the vibration testing; all the differences were within 

measurement error and typical data scatter. Differences between pre-vibe and post-vibe perveance limits 

were in general larger than for the post-vibe and post-thermal measurements. This difference is not 

significantly larger than the measurement error and typical data scatter, and thus is not necessarily an 

indication of a hardware change related to the vibration. There was no notable difference in the EBS data 

acquired throughout the testing. 

TABLE 9.—ION OPTICS FUNCTIONAL TEST DATA 

 Operating condition Pre-vibe Post-vibe Thermal 

cycle #1 

Thermal 

cycle #2 

Thermal 

cycle #3 

Postthermal 

1.2 A, 679 V 570 585 × × × 567 

2.0 A, 1179 V 672 696 × × × 680 

Perveance limit, (V) 

3.52 A, 1800 V 861 832 × × × 837 

1.2 A, 679 V –50 –50 × × × –50 

2.0 A, 1179 V –100 –103 × × × –104 

3.52 A, 1179 V –122 –122 –129 –124 –126 –120 

Electronbackstreaming limit, 

(V) 

3.52 A, 1800 V –168 –168 –168 –170 –170 –168 

A significant amount of beam current density data were acquired during the environmental testing. 

Shown in figure 12 are data acquired during two functional test operating conditions at the standard 

45 mm distance from the ion optics face. It can be seen from the data that there is essentially no 

difference in the beam profiles recorded before and after the vibration and thermal tests. 

Results of the neutralizer characterization for all functional tests are shown in figure 13. The spot-toplume-mode transition, as defined by the presence of 5 V peak-to-peak fluctuations in the neutralizer 

keeper voltage, was at 3.3 sccm for the pre-vibe and post-thermal tests and at 3.4 sccm for the post-vibe 

test. The difference in this result is not considered to be significant. 

4x10-3

3

2

1

0

Beam Current Density (A/cm2

)

-30 -20 -10 0 10 20 30

Position (cm)

 Pre-Vibe

 Post-Vibe

 Post-Thermal

3.52 A

1800 V

2.0 A

1179 V

Figure 12.—Pre- and post-vibration beam current density 

measured 45 mm from accelerator grid. 
NASA/TM—2008-215058 19

8

6

4

2

Neutralizer Keeper Fluctuations (V

0

pk-pk)

3.0 3.5 4.0 4.5 5.0 5.5 6.0

Neutralizer Flow Rate (sccm)

 Pre-Vibe

 Post-Vibe

 Post-Thermal

Fig. 13.—Neutralizer characterization test results. 

Cathode ignition times were recorded for all engine starts during the environmental test program and 

are reported in table 10. Cathode ignition time is defined as the time elapsed from the beginning of the 

ignition procedure until a full-level plasma discharge is produced. Starts are labeled ambient if the engine 

and cathode were at ambient chamber temperatures prior to the start attempt; cold starts were those 

performed after cooling of the engine with the thermal shroud at liquid nitrogen temperatures; warm starts 

were those performed within approximately one hour after previous cathode operation; and hot restarts 

were those performed within minutes after previous cathode operation. Neutralizer ignition times showed 

a slight dependence on cathode temperature during the environmental testing. The average ignition time 

for a warm/hot neutralizer was 3.5 min, and for an ambient/cold neutralizer was 4.0 min. The discharge 

cathode ignition time showed a stronger dependence on temperature. The average ignition time for 

warm/hot cathodes was 3.9 min, whereas the ambient/cold start times averaged 5.2 min. There did not 

appear to be any effect on ignition time due to the vibration or thermal-vacuum tests from the limited data 

available. Due to the longer discharge cathode ignition times and appreciable scatter in this and the 

previously reported testing (ref. 14), discharge cathode ignition time findings will require further analysis 

with regard to life validation and system-level impact. 

TABLE 10.—CATHODE IGNITION TIMES 

Test Discharge cathode 

ignition time, 

(min) 

Neutralizer 

ignition time, 

(min) 

Notes 

Pre-vibe 4.6 3.6 Warm start 

performance 6.0 5.0 Ambient start 

4.9 3.6 Warm start 

3.5 3.5 Hot restart 

Post-vibe 

performance 

5.1 4.2 Ambient start 

Thermal cycle #1 5.9 4.2 Cold start 

3.5 3.5 Hot restart 

Thermal cycle #2 4.5 3.5 Cold start 

3.5 3.5 Hot restart 

Thermal cycle #3 6.1 3.7 Cold start 

3.5 3.5 Hot restart 

Post-thermal 

performance 

3.8 3.6 Ambient start 
NASA/TM—2008-215058 20

H. Final Inspection 

After completion of the post-thermal functional testing the engine was removed from the vacuum 

facility and given a complete physical and electrical inspection. There were no significant findings as a 

part of these inspections. 

IV. Conclusion 

Environmental testing of the NEXT PM1R ion engine and gimbal assembly was completed according 

to NEXT project requirements, and both were found to be well-designed against environmental 

requirements based on the results reported herein. 

Random vibration testing was performed and successfully completed at levels of 10.0 Grms in three 

axes at 2 min per axis for the PM1R thruster and breadboard gimbal assembly. Although contamination 

was observed inside the thruster after completion of vibration testing, it was linked to sources identified 

prior to this test program that could not be completely removed in this hardware. Engineering solutions 

are available and contamination is not expected to be an issue in a flight-model hardware design. Pre- and 

post-test sine surveys showed some significant differences indicating some changes to the test hardware, 

likely a slight hardware settling at the thruster capture interface. Similar types of sine shifts were observed 

in vibration testing with this gimbal and with the PM1 engine and a thruster mass model, and are 

presently under investigation. Gimbal functional tests performed prior to and after the vibration test were 

successful with nearly identical performance in each instance. 

The thruster was successfully subjected to three thermal cycles of temperature range –120 to 215 °C. 

There were no issues found with thruster startup at cold temperatures, operation during rapid temperature 

rates of increase, operation at hot temperatures, or hot thruster restarts. 

Functional tests of the engine and the gimbal were performed prior to and after both vibration and 

thermal vacuum testing, and thruster performance was observed to be nominal throughout the test 

program. There were noticeable changes in some thruster operational parameters, likely related to modest 

differences in tank pressure between different test phases, but no indication of a major change in thruster 

performance or operation resulting from exposure to the vibration and thermal environments. Similarly, 

there was no major change in ion optics performance or beam current density data. Although neutralizer 

cathode ignition times were consistent and showed little dependence on temperature, the discharge 

cathode ignition times showed greater scatter and a stronger dependence on temperature. 

No new findings were generated as a result of the reworked PM1R environmental testing. Most of the 

findings of the previous PM1 engine testing were successfully resolved, e.g., cathode heater termination 

structural integrity and loose fasteners. The remaining issues, e.g., contamination, appear to have 

straightforward engineering solutions and should be resolved without difficulty prior to a hardware 

qualification program. 

References 

1. M.J. Patterson and S.W. Benson, “NEXT Ion Propulsion System Development Status and 

Performance,” AIAA Paper 2007–5199, 43rd Joint Propulsion Conference, Cincinnati, OH, July 

2007. 

2. D. Oh, S. Benson, K. Witzberger, and M. Cupples, “Deep Space Mission Applications for NEXT: 

NASA’s Evolutionary Xenon Thruster,” AIAA–2004–3806, 40th Joint Propulsion Conference, Fort 

Lauderdale, FL, July 2004. 

3. J.R. Brophy, “NASA’s Deep Space 1 Ion Engine,” Review of Scientific Instruments, vol. 73, no. 2, 

Feb. 2002, pp. 1071–1078. 

4. J.R. Brophy, et al., “Development and Testing of the Dawn Ion Propulsion System,” AIAA–2006–

4319, 42nd Joint Propulsion Conference, Sacramento, CA, July 2006. 
NASA/TM—2008-215058 21

5. G.C. Soulas, M.T. Domonkos, and M.J. Patterson, “Performance Evaluation of the NEXT Ion 

Engine,” AIAA–2003–5278, 39th Joint Propulsion Conference, Huntsville, AL, July 2003. 

6. G.C. Soulas, H. Kamhawi, M.J. Patterson, M.A. Britton, and M.M. Frandina, “NEXT Ion Engine 

2000 Hour Wear Test Results,” AIAA–2004–3791, 40th Joint Propulsion Conference, Fort 

Lauderdale, FL, July 2004. 

7. M.J. Patterson, L.R. Piñero, R. Aadland, and D. Komm, “NEXT Ion Propulsion System: SingleString Integration Test Results,” JANNAF Conference, May 2004. 

8. D.A. Herman, G.C. Soulas, and M.J. Patterson, “NEXT Long-Duration Test after 11,570 h and 

237 kg of Xenon Processed,” IEPC–2007–033, to be presented at the 30th International Electric 

Propulsion Conference, Florence, Italy, Sept. 17–20, 2007. 

9. M.J. Patterson, J. Foster, H. McEwen, E. Pencil, J. Van Noord, and D. Herman, “NEXT MultiThruster Array Test—Engineering Demonstration,” AIAA–2006–5180, 42nd Joint Propulsion 

Conference, Sacramento, CA, July 2006. 

10. W.A. Hoskins, F.C. Wilson, J. Polaha, L. Talerico, M.J. Patterson, G.C. Soulas, and J. Sovey, 

“Development of a Prototype Model Ion Thruster for the NEXT System,” AIAA–2004–4111, 40th 

Joint Propulsion Conference, Fort Lauderdale, FL, July 2004. 

11. D.A. Herman, G.C. Soulas, and M.J. Patterson, “Performance Evaluation of the Prototype-Model 

NEXT Ion Thruster,” AIAA–2007–5212, 43rd Joint Propulsion Conference, Cincinnati, OH, July 

2007. 

12. J.R. Anderson, J.S. Snyder, J. Van Noord, and G.C. Soulas, “Thermal Development Test of the 

NEXT PM1 Ion Engine,” AIAA–2007–5217, 43rd Joint Propulsion Conference, Cincinnati, OH, July 

2007. 

13. J.L. Van Noord, “NEXT Ion Thruster Thermal Model,” AIAA–2007–5218, 43rd Joint Propulsion 

Conference, Cincinnati, OH, July 2007. 

14. J.S. Snyder, J.R. Anderson, J. Van Noord, and G.C. Soulas, “Environmental Testing of the NEXT 

PM1 Ion Engine,” AIAA–2007–5275, 43rd Joint Propulsion Conference, Cincinnati, OH, July 2007. 

15. W.A. Hoskins, R.S. Aadland, N.J. Meckel, L.A. Talerico, and J.M. Monheiser, “NEXT Ion 

Propulsion System Production Readiness,” AIAA–2007–5856, 43rd Joint Propulsion Conference, 

Cincinnati, OH, July 2007. 

16. D.A. Vaughan, “Gimbal Development for the NEXT Ion Propulsion System,” AIAA–2005–3865, 

41st Joint Propulsion Conference, Tucson, AZ, July 2005. 

17. J.S. Snyder, M.R. O’Connell, J.P. Fernandez, G. Wang, R.S. McNabb, Jr., and D. Crumb, “Vibration 

Test of a Breadboard Gimbal for the NEXT Ion Engine, AIAA–2006–4665, 42nd Joint Propulsion 

Conference, Sacramento, CA, July 2006. 

18. NASA-HDBK-7004B, “Force Limited Vibration Testing,” NASA Technical Handbook, January 31, 

2003. 

19. S.W. Benson, M.J. Patterson, D.A. Vaughan, A.C. Wilson, and B.R. Wong, “NASA’s Evolutionary 

Xenon Thruster (NEXT) Phase 2 Development Status,” AIAA–2005–4070, 41st Joint Propulsion 

Conference, Tucson, AZ, July 2005. 

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

Space 1 Mission,” AIAA 98-3327, 34th Joint Propulsion Conference, Cleveland, OH, July 1998.

21. D.M. Goebel, R.E. Wirz, and I. Katz, “Analytical Ion Thruster Discharge Performance Model, 

AIAA–2006–4486, 42nd Joint Propulsion Conference, Sacramento, CA, July 2006. 
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Environmental Testing of the NEXT PM1R Ion Engine 

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6. AUTHOR(S)

Snyder, John, S.; Anderson, John, R.; Van Noord, Jonathan, L.; Soulas, George, C. 

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WBS 346620.04.05.03.11 

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National Aeronautics and Space Administration 

John H. Glenn Research Center at Lewis Field 

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This publication is available from the NASA Center for AeroSpace Information, 301-621-0390

13. SUPPLEMENTARY NOTES

14. ABSTRACT

The NEXT propulsion system is an advanced ion propulsion system presently under development that is oriented towards robotic 

exploration of the solar system using solar electric power. The subsystem includes an ion engine, power processing unit, feed system 

components, and thruster gimbal. The Prototype Model engine PM1 was subjected to qualification-level environmental testing in 2006 to 

demonstrate compatibility with environments representative of anticipated mission requirements. Although the testing was largely 

successful, several issues were identified including the fragmentation of potting cement on the discharge and neutralizer cathode heater 

terminations during vibration which led to abbreviated thermal testing, and generation of particulate contamination from manufacturing 

processes and engine materials. The engine was reworked to address most of these findings, renamed PM1R, and the environmental test 

sequence was repeated. Thruster functional testing was performed before and after the vibration and thermal-vacuum tests. Random 

vibration testing, conducted with the thruster mated to the breadboard gimbal, was executed at 10.0 Grms for 2 min in each of three axes. 

Thermal-vacuum testing included three thermal cycles from -120 to 215 °C with hot engine re-starts. Thruster performance was nominal 

throughout the test program, with minor variations in a few engine operating parameters likely caused by facility effects. There were no 

significant changes in engine performance as characterized by engine operating parameters, ion optics performance measurements, and 

beam current density measurements, indicating no significant changes to the hardware as a result of the environmental testing. The NEXT 

PM1R engine and the breadboard gimbal were found to be well-designed against environmental requirements based on the results reported 

herein. The redesigned cathode heater terminations successfully survived the vibration environments. Based on the results of this test 

program and confidence in the engineering solutions available for the remaining findings of the first test program, specifically the particulate 

contamination, the hardware environmental qualification program can proceed with confidence. 

15. SUBJECT TERMS

Ion thruster; Ion engine; Ion propulsion 

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