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 PM1 Ion Engine

NASA/TM—2008-215014

January 2008

AIAA–2007–5275
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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 PM1 Ion Engine

NASA/TM—2008-215014

January 2008

AIAA–2007–5275

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Prepared for the

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sponsored by the American Institute of Aeronautics and Astronautics

Cincinnati, Ohio, July 8–11, 2007
Acknowledgments

The work described in this paper was carried out by the Jet Propulsion Laboratory (JPL), California Institute of Technology, 

under a contract with the National Aeronautics and Space Administration. The PM1 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 (GRC), with 

Scott Benson as the Project Manager and Mike Patterson as the Principal Investigator. Kevin McCormick, Daniel Herman, and 

George Williams of GRC spent a signifi cant amount of time at JPL during the testing described herein and were instrumental 

in performing this work. Andy Hoskins of Aerojet also provided support for the test. At JPL, Allison Owens and Ray 

Swindlehurst 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.

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Trade names and trademarks are used in this report for identifi cation 

only. Their usage does not constitute an offi cial endorsement, 

either expressed or implied, by the National Aeronautics and 

Space Administration.

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—2008-215014 1 

Environmental Testing of the NEXT PM1 Ion Engine 

John S. 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 to demonstrate compatibility with environments representative 

of anticipated mission requirements. 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 minutes in each of three axes. Thermal-vacuum testing included a deep cold soak of the 

engine to temperatures of –168 °C and thermal cycling from –120 to 203 °C. 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. Thruster performance was nominal 

throughout the test program, with minor variations in some engine operating parameters likely caused by facility 

effects. Discharge cathode ignition times, however, were found to have a significant dependence on engine 

temperature. 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. In general, the NEXT PM1 engine and 

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

reported herein, although there were issues uncovered during the testing. After resolution of the findings from this 

test program the hardware environmental qualification program can proceed with confidence. 

I. Introduction 

 NEXT (NASA’s Evolutionary Xenon Thruster) 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 qualificationlevel testing that is representative of the anticipated environments. 

 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 Summer 2007 (Ref. 4). Initial thruster development work was led by GRC and included 

significant performance testing (Ref. 5), a 2000-hour 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 
NASA/TM—2008-215014 2 

engine was sent to 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 formal environmental test program described in this paper was initiated. 

One of the JPL activities for the NEXT development project was development of a lightweight two-axis 

breadboard gimbal for the ion thruster. 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 twenty 

degrees of thrust angle articulation while in the operational configuration (Ref. 14). 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. The breadboard gimbal was successfully subjected to vibration 

testing with a thruster mass model in an earlier test (Ref. 15). 

The objectives of the environmental test program were as follows: 

(1) To validate the NEXT PM1 thruster and breadboard gimbal assembly to qualification-level dynamic 

environments, including: 

 (a) Demonstrate thruster functionality pre- and post-vibe 

 (b) Demonstrate gimbal functionality pre- and post-vibe 

(2) To validate the NEXT PM1 thruster to expected mission thermal environments, including standard 

qualification margins, including: 

 (a) Expose the thruster to expected mission temperatures in cycles 

(b) 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 NEXT Prototype Model 1 (PM1) 

thruster. The thruster was manufactured by Aerojet (Redmond, WA) under direction of the NASA Glenn Research 

Center. Significant changes in the PM thruster 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, are presented in Reference 11. 

Prior to the environmental testing the PM1 thruster was subjected to a thermal development test in which thruster 

internal temperatures were measured under a variety of engine operating conditions and environmental boundary 

conditions (Ref. 12). Following completion of that test program the engine was disassembled, inspected, and most of 

the thermocouples were removed. The six thermocouples that remained were one on each of the three engine gimbal 

pads, and three on the front mask. 

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 

stepping 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 with a thruster mass model under identical random vibration 

loads as to be used for the testing described herein (Ref. 15). 
NASA/TM—2008-215014 3 

2. Vacuum Test Facility 

The thruster test portions of the PM1 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 nine 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 160,000 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 numerous vacuum cycles of the testing. Unfortunately, this led to modest tank 

pressure variations as the panels outgassed during thruster operation. 

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 degrees apart azimuthally with one 

end of the lamp even with the front mask of PM1. A temperature controller and SCR power controller were used for 

tight process control during temperature ramp and hold. A photograph of the thruster installed in the thermal shroud 

with the heat lamps operating is shown in Fig. 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 B. Gimbal pad B was located at the 

10 o’clock position as viewed in Fig. 1; pads A and C were located at the 6 and 2 o’clock positions, respectively. 

The thermal shroud and facility were equipped with many 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. Cathode ignitor 

circuits were incorporated into the power supply rack and 

used for cathode ignition after the vibration test; prior to 

that the cathodes had been ignited by applying steady 

voltage to the keepers. Grid recycling was handled by a 

custom recycle circuit. Pressure, flow, and electrical 

instrumentation were calibrated prior to the thermal 

development test which was performed immediately 

before the environmental 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. 

Fig. 1.—Test setup for functional and thermal 

vacuum testing. 
NASA/TM—2008-215014 4 

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 

Fig. 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 uniaxial response 

accelerometers. All accelerometers were bonded to 

Kapton (DuPont) 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 were 

mounted on the thruster directly below the interface with 

the gimbal bracket. The remaining two tri-axials 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. Z-axis, 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 

large 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. 

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. Pre-vibration 

functional testing was also the first testing after 

disassembly of the thruster and removal of the 

significant thruster thermocouple instrumentation 

following the Thermal Development Test. Thruster 

functional testing was repeated after vibration testing 

and thermal-vacuum testing. 

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 

Fig. 2.—Test setup for Z-axis vibration.
NASA/TM—2008-215014 5 

Testing of the NEXT PM1 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. 

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 –10.5 to –11.5 V for this testing. 

For the pre-vibration functional testing the discharge and neutralizer cathodes were started by application of a 

constant 150 V to the keepers until ignition, at which point the power supplies automatically went into currentlimited mode at the current set points. For subsequent functional testing, dedicated cathode ignitor circuits were used 

to apply rapid high-voltage pulses for plasma discharge ignition. 

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-of-change 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 Fig. 3. The gimbal was driven via a custom software 

program and dedicated motor control hardware. 

Table 2.—Thruster Controlled Parameters for Functional Testing

Control Parameter Discharge-Only FTP 1 FTP 2 FTP 3 FTP4 

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 
NASA/TM—2008-215014 6 

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). 

3. Vibration Test 

The random vibration specification for the NEXT 

gimbal and thruster assembly was based on Delta-class 

launch vehicles and typical locations of electric propulsion 

hardware on spacecraft. The qualification-level 

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 two minutes per axis. 

Sine surveys were performed before and after each fulllevel 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. 

Vibration testing of the NEXT gimbal also incorporated the practice of force-limited 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. 16). 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 expander, the 

facility was reconfigured for lateral axis testing, and the test 

assembly attached to the slip plate. 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. A –

6 dB intermediate-level run was performed before the two-minute 

random vibration test at the full load, which was followed by a 

post-random 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. 17). 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 38o 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). 

Fig. 3.—Gimbal functional test.

Table 3.—Qualification-Level Random 

Vibration Requirements.

Frequency (Hz) Level 

20 Hz 0.04 g2/Hz 

20 – 50 Hz +3 dB/octave 

50-600 Hz 0.1 g2/Hz 

600 – 2,000 Hz -6 dB/Octave 

2000 Hz 0.009 g2/Hz 

Overall 10.0 Grms

Duration: 2 min./axis, 3 orthogonal axes 
NASA/TM—2008-215014 7 

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. 18). 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 NEXT 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 PM1 thermal model 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. 

Preliminary results of that work indicated that a gimbal pad temperature of 183 °C would be expected, yielding a 

thermal-vacuum test maximum temperature of 203 °C including the standard flight temperature margin. (Note that 

recent refinements in the model have suggested maximum temperature of 187 °C under these conditions (Ref. 13), 

indicating a higher maximum temperature may be required for future development testing.) 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. 18). 

The plan for the thermal vacuum test was to perform three identical thermal cycles (as will be shown later, only 

two cycles were actually performed). The cycles consisted of: 

(1) Two-hour 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 203 °C. 

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

(5) Two hours 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 Fig. 4. In this test profile, three cold starts and three hot starts were 

planned. The planned total cumulative cold soak time was 6 hours and total cumulative hot operation time was 

12.5 hours. 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 90 minutes of operation at each of the two beam extraction operating points, electron backstreaming 

data were acquired. It was also planned to acquire beam current density data with the Faraday probes, but after the 

first scan during the first thermal cycle the stage system malfunctioned and no more data could be acquired. 

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150

100

50

0

-50

-100

-150

-200

Reference Temperature (oC)

0 10 20 30 40 50 60 70 80 90

Time (hours)

 Thruster ON

 Thruster OFF

Times are approximate

Fig. 4.—Thermal vacuum test plan temperature profile. 
NASA/TM—2008-215014 8 

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 

Performance Acceptance Test (PAT) conducted at GRC (Ref. 11) and the pre-vibration functional test along with 

ion optics performance measurements. The modest differences in the discharge voltage and accelerator grid current 

are almost certainly due to differences in the tank pressures between the two facilities in which these tests were run 

(about a factor of 3.5 lower in the GRC facility at full power). 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. Apart from these parameters, there was no significant 

difference observed in the engine performance data. Ion optics performance margins were the same within 

measurement error with the exception of a notable difference in perveance margin at the 2.0 A, 1179 V operating 

condition. Though this numerical comparison appears to indicate a difference somewhat larger than the nominal 

measurement error, direct comparison of the data used to infer the perveance limit suggest that there were no 

significant differences in optics performance at this condition and that the numerical differences were due to 

relatively large voltage steps used when acquiring data near the limit in one case. 

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 ±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. 

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 Fig. 2. Initial tests 

and then the two-minute 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 

Fig. 5. Significant notches were observed at 100, 750, and 1150 Hz. 

Table 4.—Comparison of Engine Performance Data from Initial Performance 

Acceptance Test (PAT) (Ref. 11) and Pre-Vibration Functional Test

 3.52 A, 1800 V 2.0 A, 1179 V 

Parameter Initial 

PAT Pre-Vibe Initial 

PAT Pre-Vibe 

Discharge Current (A) 18.4 18.1 13.8 13.9 

Discharge Voltage (V) 24.3 23.2 24.7 24.5 

Cathode Keeper Voltage (V) 5.3 5.4 3.9 3.9 

Accelerator Grid Current (mA) 14.5 20.8 6.1 7.2 

Neutralizer Keeper Voltage (V) 12.1 12.0 13.7 13.3 

Power (W) 6860 6860 2750 2760 

Thrust (mN) 237 238 108 109 

Specific Impulse (sec) 4190 4190 3490 3500 

Total Efficiency 0.710 0.712 0.675 0.677 

perveance Margin (V) 1150 1156 720 686 

Electron Backstreaming Margin (V) 43 49 97 101 
NASA/TM—2008-215014 9 

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° 17° 

 Level A, Max B, Min C 10 -1° -17° 

Stowed, Latched 0 0° 0° 

Inspection of the test article immediately after 

the full-level test revealed a large amount of debris 

and particulate contamination in the bottom of the 

thruster plasma screen cone. This debris was 

eventually traced to fragmentation of potting 

cement on the discharge cathode heater electrical 

termination. Particulate dust was also observed 

within the plasma screen, on the discharge 

chamber stiffening ring near the gimbal pads. 

Electrical inspection of the engine after the 

Z-axis testing revealed significant changes in the 

discharge cathode heater resistance. The heater 

resistance increased from its nominal value, as 

measured prior to the full test, by a factor of 

seventeen, well out of specification. No significant 

changes were noted in the neutralizer heater 

resistance. The discharge cathode heater resistance 

continued to vary appreciably throughout the 

vibration testing, and was later correlated to the 

potting cement fragmentation. 

Comparison of the pre- and post-vibe sine 

surveys showed significant changes on some 

locations of the test assembly near the fundamental 

mode. One of the largest changes, shown in Fig. 6, 

was observed on the gimbal thruster bracket at leg 

C. Two different effects are observed near 100 Hz, 

a small change in frequency and a significant 

damping of the response. The observed frequency 

shift of roughly 5 percent is within what is 

nominally observed in run-to-run comparison. The 

change in damping is not necessarily an indicator 

for a serious problem but may be, for example, a 

fastener loosening somewhat. The two issues are 

not necessarily the same and could result from 

different effects. Similar changes were observed in 

the accelerometer mounted on the thruster gimbal 

pad, on the other side of the gimbal-thruster 

interface from the data of Fig. 6. 

 

0.001

0.01

0.1

1

Power Spectral Density (g2/Hz)

10 2 3 4 5 6 7 8100 2 3 4 5 6 7 81000 2

Frequency (Hz)

 Specification (10.0 Grms)

 Control (9.3 Grms)

Fig. 5.—Comparison of random vibration specification 

with force-limited test inputs, Z-axis excitation.

10-4

10-3

10-2

10-1

100

101

Accelerometer Response (G's)

10 2 3 4 5 6 7 8100 2 3 4 5 6 7 81000 2

Frequency (Hz)

 Pre-Vibe

 Post-Vibe

Accelerometer A9Z

Fig. 6.—Comparison of sine survey results before and 

after full-level Z-axis random vibration, gimbal 

thruster bracket leg C. 
NASA/TM—2008-215014 10 

In contrast to the changes observed at leg C, 

there was little change observed in the data from 

accelerometers location near the other two gimbal 

thruster brackets on legs A and B. Moderate 

differences in sine surveys were seen in the other 

accelerometers located on the engine. There was 

little if any change in the sine survey data for the 

accelerometers on the vibration test mounts, as 

shown in the representative data of Fig. 7. These 

results collectively indicate that changes occurred 

in the test article near the thruster/gimbal interface 

at leg C and not at the structural bolting interface 

between the vibration fixturing and the vibe table. 

Taken by themselves, the sine survey shifts are not 

especially troubling but when combined with the 

other observed thruster anomalies they required 

further review. 

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. After removal from the vibration table, a torque check was 

performed on the test assembly fasteners. Slight loosening was observed on the stabilizer arm fasteners of gimbal 

legs B and C. Also observed was a slight rotation in the thruster pin at leg A. The fasteners joining the thruster and 

gimbal had not loosened. 

Additional review of the sine survey data after conclusion of the vibration testing suggested that hardware 

changes occurred in the main load path near gimbal pad C. Changes in sine survey data for accelerometers at gimbal 

pad C were similar, and were much larger than those observed for the accelerometers at pads A and B. It is likely 

that the source of the shift was between the accelerometer on the gimbal thruster bracket and the gimbal mounting 

plate, but it is also possible that the source was internal to the thruster in the load path of the gimbal bracket. During 

thruster disassembly and inspection after the completion of the environmental test program, however, there was no 

evidence of a cause for a sine survey shift in the thruster hardware. At this point, investigation of the sine survey 

shift was concluded with the most likely cause a slight settling in the hardware at the leg C thruster gimbal bracket. 

2. Lateral Axis Testing 

Dynamic testing continued with the Y-axis testing. The test setup is shown in the photograph of Fig. 8. The 

Y-axis testing proceeded nominally until an intermediate-level random vibration test in which a loose fastener on the 

thruster plasma screen was observed and the test was aborted. The torque on all plasma screen fasteners was 

checked after this test and three were found to 

have loosened noticeably. It is likely that some of 

this loosening occurred during the Z-axis testing 

but was not observed until this time. After 

tightening the loose fasteners and repeating the 

pre-random sine survey, the two-minute full 

random vibration test was completed without 

incident. 

After completion of the Y-axis testing the test 

assembly shown in Fig. 8 was rotated 90° to 

perform the X-axis random vibration testing. The 

X-axis test sequence including the two-minute full 

random vibration test was completed without 

incident. The vibration test inputs for both X and 

Y excitation are shown in Fig. 9. Significant 

notches in the control acceleration spectrum were 

observed at 80 and 390 Hz for the Y-axis 

excitation, and at 82, 380, and 1150 Hz for X-axis 

excitation. 

10-4

10-3

10-2

10-1

100

101

Accelerometer Response (G's)

10 2 3 4 5 6 7 8100 2 3 4 5 6 7 81000 2

Frequency (Hz)

 Pre-Vibe

 Post-Vibe

Accelerometer A18Z

Fig. 7.—Comparison of sine survey results before and 

after full-level Z-axis random vibration, vibe test 

mount C, near motor. 

Fig. 8.—Test setup for Y-axis vibration.
NASA/TM—2008-215014 11 

Physical inspection of the test assembly after 

the Y-axis testing indicated some slight movement 

in the thruster pin at gimbal leg B, with no other 

significant findings. Likewise, slight movement of 

thruster pins at gimbal legs B and C was observed 

after X-axis testing with no other significant 

findings. 

Electrical inspection of the test assembly after 

the Y-axis testing revealed that, in addition to 

variations in the discharge cathode heater 

resistance, the neutralizer heater resistance had 

increased from the nominal value measured before 

the test to open-circuit as measured by a digital 

multimeter, indicating a physical heater failure 

caused by Y-axis vibration testing. Tapping of the 

neutralizer housing with a plastic tool caused 

considerable resistance changes with each tap, 

suggesting a loose electrical connection. After 

completion of the X-axis vibration testing, both 

discharge and neutralizer heaters were found to be 

open circuit. 

Comparison of the Y-axis pre- and post-vibe 

sine surveys showed minor changes on the test 

assembly near the fundamental mode in the 

frequency range of 50-100 Hz. Typical data are 

shown in Fig. 10; similar differences were 

observed throughout the other accelerometers on 

the test article. Minor changes near the 

fundamental mode were also observed after X-axis 

testing among the accelerometers located on the 

test article. These sine survey differences are 

considered to be quite minor and are not an 

indication of a significant change in the test 

hardware. 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. 

Physical inspection of the test article before removal from the vibration table following X-axis testing showed no 

evidence of loosening of any fasteners on the gimbal stabilizer arms, thruster plasma screen, or the twelve fasteners 

joining the thruster and gimbal. Torque on the twelve bolts fastening the vibration test mounts to the vibration test 

plate through the force transducers was checked and found to be the same as the installation torque (i.e. no loosening 

was observed). 

3. Post-Vibration Inspection Results 

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

testing. There were no significant additional findings as a part this inspection, though contamination and cathode 

heater resistance findings were followed up in greater detail. In many instances the amount of contamination 

generated was in excess of typical spacecraft cleanliness requirements. 

The interior of the thruster discharge chamber was visually inspected through the ion optics and a significant 

amount of particulate contamination was observed. The material appeared to be a white fine-granular powder 

concentrated near the discharge cathode on either side of the cathode magnet ring, as shown in Fig. 11. The material 

was later sampled and shown to be residue from a mechanical fabrication treatment applied to the discharge 

chamber during engine fabrication. 

0.001

0.01

0.1

1

Power Spectral Density (g2/Hz)

10 2 3 4 5 6 7 8100 2 3 4 5 6 7 81000 2

Frequency (Hz)

 Specification (10.0 Grms)

 X-Axis Control (9.7 Grms)

 Y-Axis Control (9.8 Grms)

Fig. 9.—Comparison of random vibration specification 

with force-limited test inputs, lateral axis excitation. 

10-3

10-2

10-1

100

Accelerometer Response (G's)

10 2 3 4 5 6 7 8100 2 3 4 5 6 7 81000 2

Frequency (Hz)

 Pre-Vibe

 Post-Vibe

Accelerometer A7R

Fig. 10.—Comparison of sine survey results before and 

after full-level Y-axis random vibration, gimbal 

thruster bracket leg A.
NASA/TM—2008-215014 12 

Internal surfaces of the thruster were inspected by 

removing a portion of the cylindrical plasma screen. Most 

of the particulate dust that was observed after Z-axis 

testing appeared to have been shaken off during the 

subsequent lateral axis testing, but some particulates 

remained on the discharge chamber stiffening ring 

surfaces. Visible particulate material was sampled and the 

composition was determined to be a mixture of metallic 

particles, residue from fabrication processes, and loose 

fibers. The fibers were positively identified as a match to 

the fiberglass sleeving on the internal thruster wiring. 

Metallic residue was traced to debris from nut plate 

galling. 

The cover plate of the neutralizer assembly was 

removed as a part of the post-test inspection to investigate 

the cause of the heater resistance variations. Debris of 

various sizes was observed at the bottom of the housing 

not unlike the debris seen at the bottom of the plasma 

screen during Z-axis vibration testing. The heater 

termination area showed significant loss of cement potting 

compound, clearly the source of the observed debris. This 

indicated, and a full inspection later confirmed, that the 

debris found in the plasma screen cone during Z-axis vibration resulted from freed potting cement on the discharge 

cathode assembly. Inspection of the neutralizer low-voltage propellant isolator revealed no visible cracks or other 

anomalies. Inspection through the housing downstream aperture toward the cathode and keeper orifice plates 

revealed nothing of note, but visual access was limited. 

D. Post-Vibration Gimbal Functional Test 

Gimbal functional testing was repeated in the same manner that it had been performed in the pre-vibration 

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 radial inclination in the “Level A, Max B, Min C” step which 

was –1° in the pre-vibe test and 0° in the post-vibe 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. 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. 

Cathode heater continuity was achieved by applying short-duration pulses of moderate voltage and current to the 

heaters, after which nominal resistance was restored. 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 

Immediately upon conclusion of the post-vibration engine performance testing, LN2 flow to the thermal shroud 

was initiated and the shroud door was closed to begin an extended cold soak of the engine. After 114 hours of 

soaking the engine reached temperatures of –168.1 to –168.6 °C on the gimbal pads and –167.4 to 167.5 °C on the 

front mask. The engine was demonstrated to start nominally after this soak and was operated in discharge-only mode 

for approximately thirty minutes before it was shut down. Following this, a non-operational thermal cycle was 

performed. The thermal shroud heat lamps were turned on and the reference temperature location (gimbal pad B) 

was brought from 66 °C to the control temperature of 203 °C in about 100 minutes, where it was held for 

152 minutes before turning off the lamps. 

Fig. 11. Discharge Chamber Particulate 

Contamination (photographed through 

ion optics). 
NASA/TM—2008-215014 13 

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

ignition process was initiated two 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 203 °C control temperature in 46 minutes. Thruster operation was held 

steady at full power (3.52 A, 1800 V) and at the control temperature for 134 minutes before transitioning to the 

lower power point (3.52 A, 1179 V). After 125 minutes 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 were allowed to cool for the next thermal cycle. Performance data obtained during the thermal cycles 

are discussed in a subsequent section. The temperature history for all of the thermal-vacuum testing is shown in 

Fig. 12. Note that the coolest of the three gimbal pads was chosen for the control thermocouple, hence the 

temperatures of the other two gimbal pads were higher than the 203 °C control temperature as seen in the figure. 

The second thermal cycle was performed on the subsequent day. The cathode ignition process was initiated two 

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 turned on and the temperature reference location ramped to the 

203 °C control temperature in 48 minutes. Thruster operation was held steady at full power and the control 

temperature for 133 minutes before transitioning to the lower power. After 120 minutes 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 then at full power before finally 

turning the engine off. The engine and thermal shroud were allowed to cool for the next thermal cycle. 

On the third day of thermal-vacuum testing the neutralizer heater failed open-circuit when attempting to perform 

the cold start. After application of environmental heating (i.e. discharge-only operation and low-level lamp heating) 

for 38 minutes the neutralizer heater was able to operate. At this point it was decided to skip the third thermal cycle 

and proceed to functional testing. 

G. Thruster Functional Testing 

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

test program. Despite the difficulties in starting the neutralizer heater at the end of the thermal-vacuum testing, the 

engine performed nominally during all of the final performance testing. Non-controlled and derived engine 

performance parameters for the lower-power 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. 

Discharge chamber parameters and accelerator grid current were likely influenced by the tank pressure differences 

of up to 10 percent between tests, caused by graphite panel gas absorption/desorption during the frequent 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. 19). Neutral density 

near the accelerator grid is also a significant factor in charge-exchange ion production and accelerator grid current. 

250

200

150

100

50

0

-50

-100

-150

-200

Temperature (°C)

11/13/2006 11/14/2006 11/15/2006 11/16/2006

Runtime

 Front Mask Thermocouples

 Gimbal Pad Thermocouples

Cycle #1 Cycle #2 Functional

Fig. 12.—Thermal vacuum testing temperature profile.
NASA/TM—2008-215014 14 

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. 

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

performance from the two 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 higher measure tank pressures during the thermal cycle testing at elevated temperatures. 

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. Comparison of 

EBS data with the 3.52 A, 1179 V operating condition in the pre-vibe test was inconclusive because of the difficulty 

in getting a good measurement with the frequent engine recycling during the measurement. 

Differences between post-vibe and post-thermal perveance limits were in general larger than for the pre- and 

post-vibe measurements. This difference is not significantly larger than the measurement error, and thus is not 

necessarily an indication of a hardware change related to the thermal-vacuum test. There was no notable difference 

in the EBS data for the post-vibe, thermal-vacuum, and post-thermal tests. 

Table 6.—Engine Performance for the Lower-Power Functional Test Points

 1.2 A, 679 V 2.0 A, 1179 V 

 Pre-Vibe PostVibe 

PostThermal Pre-Vibe Post-Vibe PostThermal 

Discharge Current (A) 10.2 10.4 9.9 13.9 13.9 13.7 

Discharge Voltage (V) 26.6 25.8 26.2 24.5 23.7 23.9 

Cathode Keeper Voltage (V) 4.0 3.7 4.0 3.9 3.9 3.9 

Accelerator Grid Current (mA) 3.7 3.8 3.9 7.2 7.5 7.8 

Neutralizer Keeper Voltage (V) 13.5 14.3 13.9 13.3 13.8 13.5 

Power (W) 1130 1130 1120 2760 2740 2740 

Tank Pressure (Torr Xe) 1.9E-06 1.7E-06 2.1E-06 3.0E-06 2.7E-06 3.2E-06 

Thrust (mN) 49.2 49.2 49.1 109 109 108 

Specific Impulse (sec) 2440 2440 2440 3500 3480 3480 

Total Efficiency 0.521 0.522 0.526 0.677 0.676 0.676 

Beam Ion Energy Cost (eV/ion) 227 224 216 169 165 163 

Table 7.—Engine Performance for the 3.52 A, 1179 V Operating Condition 

 Pre-Vibe PostVibe 

Thermal 

Cycle #1 

Thermal 

Cycle #2 

PostThermal 

Discharge Current (A) 19.5 19.3 19.2 19.3 19.5 

Discharge Voltage (V) 23.9 24.1 23.4 23.4 23.5 

Cathode Keeper Voltage (V) 5.5 5.5 5.5 5.4 5.3 

Accelerator Grid Current (mA) 20.7 18.4 29.5 26.2 20.7 

Neutralizer Keeper Voltage (V) 12.0 11.9 12.2 12.0 12.0 

Power (W) 4660 4660 4670 4660 4660 

Tank Pressure (Torr Xe) 5.1E-06 4.9E-06 6.3E-06 5.9E-06 5.6E-06 

Thrust (mN) 191 191 191 191 191 

Specific Impulse (sec) 3380 3380 3380 3380 3380 

Total Efficiency 0.679 0.679 0.679 0.680 0.679 

Beam Ion Energy Cost (eV/ion) 132 132 128 128 130 
NASA/TM—2008-215014 15 

A significant amount of beam current density 

data were acquired during the pre- and postvibration functional testing. Unfortunately, the 

mechanical stage used to articulate the Faraday 

probes failed at the beginning of the thermal 

vacuum testing and could not be used to collect 

data at the standard test conditions for the 

remainder of the testing. A subset of the available 

data is presented here to compare the functional 

results before and after vibration testing. Shown in 

Fig. 13 are data acquired during two functional test 

operating conditions at the standard 45 mm 

distance from the ion optics face. The slight 

differences between data collected before and after 

the vibration test are typical of those seen at the 

other operating conditions and distances. To 

investigate if a significant shift had occurred as a 

result of vibration testing, the probe data were 

examined in greater detail and it was determined that the differences observed were not significantly larger than the 

errors associated with collecting and analyzing the data. This conclusion is consistent with the similar results 

obtained for ion optics measurements. 

Table 9.—Ion Optics Functional Test Data

Operating 

Condition Pre-Vibe PostVibe 

Thermal 

Cycle #1 

Thermal 

Cycle #2 

PostThermal 

1.2 A, 679 V 578 584 × × 568 

2.0 A, 1179 V 693 687 × × 695 

3.52 A, 1179 V 861 863 × × 851 

Perveance 

Limit (V) 

3.52 A, 1800 V 855 859 × × 844 

1.2 A, 679 V –49 –49 × × –47 

2.0 A, 1179 V –99 –97 × × –98 

3.52 A, 1179 V –137* –117 –119 –119 –118 

ElectronBackstreaming 

Limit (V) 3.52 A, 1800 V –161 –165 –161 –166 –164 

*NOTE: there was significant engine recycling during this measurement that likely affected the result. 

4x10-3

3

2

1

0

Beam Current Density (A/cm2

)

-30 -20 -10 0 10 20 30

Position (cm)

1800 V 3.52 A

 Pre-Vibe

 Post-Vibe

1179 V 2.0A

 Pre-Vibe

 Post-Vibe

Fig. 13.—Pre- and post-vibration beam current density 

measured 45 mm from accelerator grid.

Table 8.—Engine Performance for the 3.52 A, 1800 V Operating Condition 

 Pre-Vibe PostVibe 

Thermal 

Cycle #1 

Thermal 

Cycle #2 

PostThermal 

Discharge Current (A) 18.1 18.0 18.0 18.1 17.8 

Discharge Voltage (V) 23.2 23.4 23.0 22.9 23.5 

Cathode Keeper Voltage (V) 5.4 5.4 5.4 5.4 5.5 

Accelerator Grid Current (mA) 20.8 18.2 25.8 21.7 19.8 

Neutralizer Keeper Voltage (V) 12.0 12.0 12.2 12.0 12.1 

Power (W) 6860 6840 6840 6830 6820 

Tank Pressure (Torr Xe) 5.5E-06 5.2E-06 6.2E-06 5.9E-06 5.5E-06 

Thrust (mN) 238 237 237 237 236 

Specific Impulse (sec) 4190 4190 4170 4170 4180 

Total Efficiency 0.712 0.712 0.708 0.709 0.710 

Beam Ion Energy Cost (eV/ion) 119 120 117 118 119 
NASA/TM—2008-215014 16 

Results of the neutralizer characterization for 

the post-vibe and post-thermal functional tests are 

shown in Fig. 14. Results for the pre-vibration 

testing are not shown because interpretation of the 

data was difficult due to experimental 

complications. The spot-to-plume-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 both tests in Fig. 14. This value is 

the same as that measured in the PM1 performance 

evaluation conducted prior to the thermal 

development test (Ref. 11). 

Cathode ignition times were recorded for all 

engine starts during the environmental test 

program, and a subset of those results are 

summarized in Table 10. Cathode ignition time is 

defined as the time elapsed from the beginning of 

the ignition procedure until a plasma discharge is 

produced. Starts are labeled ambient if the engine 

and cathode were a 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 minutes, 

and for an ambient/cold neutralizer was 4.3 minutes. In contrast, the discharge cathode ignition time showed a 

strong dependence on temperature. The average ignition time for warm/hot cathodes was a nominal 3.6 minutes, 

whereas the ambient/cold start times averaged 6.5 minutes. There did not appear to be any effect on ignition time 

due to the vibration or thermal-vacuum tests from the limited data available. Discharge cathode ignition time 

findings will require further analysis with regard to life validation and system-level impact before it is determined if 

a hardware change is necessary. 

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

Neutrazlier Flow Rate (sccm)

 Post-Vibe

 Post-Thermal

Fig. 14.—Neutralizer characterization test results. 

Table 10.—Cathode Ignition Times 

Test 

Discharge 

Cathode 

Ignition 

Time (min) 

Neutralizer 

Ignition 

Time (min) 

Notes 

8.7 4.0 Ambient start 

6.4 4.2 Ambient start Post-Vibe 

Performance 

6.9 4.2 Ambient start 

6.1 4.2 Cold start 

5.9 4.4 Cold start 

5.3 3.8 Ambient start 

3.7 3.5 Warm start 

Pre-ThermalVacuum 

Operation 

6.3 4.3 Ambient start 

Thermal 7.2 4.6 Cold start 

Cycle #1 3.5 3.5 Hot Restart 

Thermal 6.1 4.6 Cold start 

Cycle #2 3.5 3.5 Hot Restart 

5.6 × Cold start 

Post-Thermal 

Performance × 3.5 

Heater failed at first start 

attempt, later started with 

environmental heating 
NASA/TM—2008-215014 17 

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. The discharge and neutralizer cathode impedances were nominal 

and there were no other significant findings as a part of the electrical inspection. Likewise, there were no significant 

additional findings as a part of the physical inspection. It was noted, however, that the debris observed within in the 

discharge chamber (see Fig. 11) no longer had the same localized concentration, likely a result of handling the 

thruster and placing it in a horizontal orientation for testing. The neutralizer housing cover was removed for visual 

inspection which revealed no immediate cause for the heater failure during thermal-vacuum testing. 

IV. Conclusion 

Environmental testing of the NEXT PM1 ion engine and gimbal assembly was completed according to NEXT 

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

and thermal vacuum testing. 

Random vibration testing was performed at levels of 10.0 Grms in three axes at two minutes per axis for the PM1 

thruster and breadboard gimbal assembly. Testing in each axis was successfully completed, although with several 

issues uncovered. 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. 

In addition, pre- and post-test sine surveys showed some significant differences indicating some changes to the 

test hardware. The largest of those was identified as a change in the load path near gimbal leg C, likely in the 

breadboard gimbal hardware but possibly occurring within the thruster immediately next to the gimbal pad. 

Inspection of the test hardware yielded no evidence of a cause for the sine survey shift, hence the most likely cause 

was determined to be a slight hardware settling at the leg C thruster gimbal bracket. Gimbal functional tests 

performed prior to and after the vibration test were successful with nearly identical performance in each instance. 

Thermal vacuum testing began with a deep cold soak to temperatures below –160 °C on the PM1 gimbal pads 

and front mask. Following a non-operational thermal cycle, the thruster was successfully subjected to two of three 

planned thermal cycles of temperature range –120 °C to 203 °C. On the third attempt to start the thruster the 

neutralizer heater failed to pass current, most likely related to the impedance anomalies observed in vibration 

testing. Environmental heating was applied to the neutralizer and heater operation was recovered, allowing engine 

operation. Post-thermal-vacuum functional testing was then successfully performed. 

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. 

In general, the NEXT PM1 engine and the breadboard gimbal were found to be well-designed against 

environmental requirements based on the results reported herein, although there were issues uncovered during the 

testing. The minor findings (e.g. loose fasteners) should be easily resolvable. With modest redesign efforts targeted 

at the cathode heaters, and with materials and processes solutions to the contamination findings, these issues can 

also be corrected. Discharge cathode ignition time findings will require further analysis with regard to life validation 

and system-level impact before it is determined if a hardware change is necessary. After resolution of the findings 

from this test program the hardware environmental qualification program can proceed with confidence. 

References 

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

Paper 2007–5199, to be presented at the 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. 
NASA/TM—2008-215014 18 

 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. 

 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: Single-String 

Integration Test Results,” JANNAF Conference, May 2004. 

 8. D.A. Herman, G.C. Soulas, and M.J. Patterson, “Status of the NEXT Ion Thruster Long Duration Test,” AIAA–

2007–5272, to be presented at the 43rd Joint Propulsion Conference, Cincinnati, OH, July 2007. 

 9. M.J. Patterson, J. Foster, H. McEwen, E. Pencil, J. Van Noord, and D. Herman, “NEXT Multi-Thruster 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, to be presented at the 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, to be presented at the 43rd Joint Propulsion Conference, Cincinnati, OH, July 

2007. 

13. J.L. Van Noord, “NEXT Ion Thruster Thermal Model,” AIAA–2007–5218, to be presented at the 43rd Joint 

Propulsion Conference, Cincinnati, OH, July 2007. 

14. D.A. Vaughan, “Gimbal Development for the NEXT Ion Propulsion System,” AIAA–2005–3865, 41st Joint 

Propulsion Conference, Tucson, AZ, July 2005. 

15. 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. 

16. NASA–HDBK–7004B, “Force Limited Vibration Testing,” NASA Technical Handbook, January 31, 2003. 

17. 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. 

18. 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.

19. 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 PM1 Ion Engine 

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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 Prototype Model engine PM1 was subjected to qualification-level

environmental testing to demonstrate compatibility with environments representative of anticipated mission requirements. Random vibration 

testing, conducted with the thruster mated to the breadboard gimbal, was executed at 10.0 Grms for 2 minutes in each of three axes. Thermalvacuum testing included a deep cold soak of the engine to temperatures of –168 °C and thermal cycling from –120 to 

203 °C. 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. Thruster performance was nominal throughout the test program, with 

minor variations in some engine operating parameters likely caused by facility effects. In general, the NEXT PM1 engine and the

breadboard gimbal were found to be well-designed against environmental requirements based on the results reported herein. After resolution 

of the findings from this test program the hardware environmental qualification program can proceed with confidence. 

15. SUBJECT TERMS

Ion thruster; Ion engine; Ion propulsion 

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