L-3 COMMUNICATIONS ETI ELECTRIC PROPULSION OVERVIEWelectricrocket.org/IEPC/315.pdf · 2020. 1....

The 29th International Electric Propulsion Conference, Princeton University,

October 31 – November 4, 2005

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L-3 Communications ETI Electric Propulsion Overview

IEPC-2005-315 Presented at the 29th International Electric Propulsion Conference, Princeton University,


Kuei-Ru Chien*, Stephen L Hart†, William G. Tighe‡, Michael K De Pano§, Thomas A Bond** and Rafael Spears** L-3 Communications Electron Technologies, Inc., 3100 W. Lomita Blvd., Torrance, California 90509, USA

L-3 Communications Electron Technologies Inc. (ETI) is a new company (formerly Hughes/Boeing EDD) with more than 40 years experience in electric propulsion (EP). The EP Product Line is currently producing 25cm xenon ion thrusters (XIT) and power controller units (XPC) for the Boeing Satellite Systems 702 satellites. There are 100 combined 13 cm and 25 cm xenon ion thrusters in orbit with over 125,000 cumulated flight hours. Life test results of 21,000 hours have been obtained with the 13 cm xenon ion thruster and over 15,550 hours of 25 cm xenon ion thruster life testing have been completed to date. ETI is also producing 30cm xenon ion thrusters and power processor units (PPU) for the Jet Propulsion Laboratory’s Dawn spacecraft and developing the power processor unit for NASA GRC 40cm NEXT ion thruster. In this paper, we will provide an overview of our manufacturing process, test capabilities and life test results.

I. Introduction L-3 Communications Electron Technologies Inc. (ETI) is a new company (formerly Hughes/Boeing EDD) with

a long history in the development of Electric Propulsion (EP) technology. Decades of research and testing have led ETI to a comprehensive understanding of this technology and its manufacturing processes. The technology development started in the early 1960s. In 1992, ETI’s ion thruster discharge chamber successfully flew aboard the NASA Space Shuttle. Experiments conducted on the shuttle flight determined the ability to neutralize negative charge buildup on space structures. These tests were key steps in moving ion propulsion technology from the laboratory into the factory. The flight production then started in 1994. The first xenon ion propulsion system (XIPS), which featured a 13 cm active grid diameter, was launched in 1997 on the Boeing 601HP satellite. As of today, there are 100 combined 13cm and 25 cm xenon ion thrusters in the orbit with over 125,000 cumulated on hours. In this paper, we will describe the performances of the three different XIPS (ion thruster and power controller) configurations, the overall manufacturing and test processes and capabilities, and the life test results pertaining to ETI’s 13 cm and 25 cm ion thrusters.

II. ETI Flight Ion Thruster and Power Controller Performances ETI EP Product Line currently has three different XIPS configurations available1. These options allow its

customers to take advantage of the technology optimized for various spacecraft size and power ranges. At the

* Chief Scientist, Electric Propulsion Product Line, [email protected] † Project Engineer, Electric Propulsion Product Line, [email protected] ‡ Physicist/Engineer, Electric Propulsion Product Line, [email protected] § Physicist/Engineer, Electric Propulsion Product Line, [email protected] **Program Manger, Electric Propulsion Product Line, [email protected] **Vice President, Electric Propulsion Product Line, [email protected]



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present time, thirty-six 25cm thrusters and 18 XPCs are currently in production at ETI. Table 1 indicates the current status of these three product lines.

Table1. ETI Flight Ion Thruster Production Status

Thruster Program Quantity Status

13cm Boeing 601HP 60 In Orbit 25cm Boeing 702 40 In Orbit 25cm Boeing 702 32 In Production 30cm Deep Space 1 2 Mission Completed 30cm Dawn 3 In Spacecraft Integration

The performances of these three products are described below.

13cm Xenon Ion Thruster and Power Processor Unit The 13 cm EP system is used on the Boeing 601HP commercial communications satellite. It consists of four

13cm xenon ion thrusters and two power processors in a fully redundant configuration to provide north-south station-keeping, momentum dumping and eccentricity control. The production of the 13cm flight thrusters began in 1994. PAS5, the first 601HP satellite with electric propulsion was launched in 1997. Since then ETI has continued to make improvements to the design and manufacturing processes throughout the production run. Sixty flight 13 cm thrusters have been produced, launched and have achieved more than 91,000 cumulated in orbit operation hours.

The thruster consists of conventional hollow cathodes for the discharge assembly and neutralizer, a three-grid

ion optics assembly, permanent magnets, and a discharge/structure chamber. The operation temperature ranges from -110°C to + 172°C. Typical 13 cm ion thruster performance is shown in table 2.

Table2. 13cm Ion Thruster Performance

Parameter Performance

Total Input Power (W) 450 Thrust (mN) 18

Specific Impulse (s) 2350 Electrical Efficiency (%) 68

Mass Utilization Efficiency (%) 72 Beam Voltage (V) 750

Beam Current (mA) 400 The power processor unit (PPU) serves two functions – control of the thruster and interfacing with the spacecraft system. The PPU takes the raw spacecraft bus power and conditions it to the power levels needed by the thruster. Each PPU can independently operate one north-south ion thruster pair. As a system interface the PPU does the following: it provides timing and sequencing for thruster on and thruster off commands, performs fault protection to avoid damage to the thruster and any of the spacecraft components, performs grid clearing in the case of a particle being caught between two of the electrode grids and provides telemetry for the purpose of measuring thruster performance and subsystem state of health. Table 3 summarizes PPU typical performances.

Table3. 13cm PPU Performance


Total Input Power (W) 530 Bus Input Voltage (V) 49-53 PPU Efficiency (%) 86

Size (cm) 28x20x44 Mass (kg) 14.6



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25cm Xenon Ion Propulsion System (XIPS) The 25 cm EP system is used on Boeing 702 communications satellites2. It uses four 25 cm xenon ion thrusters

and two power controllers. This is the first commercial satellite to use EP exclusively for attitude control, all station keeping (for both north-south station keeping and east-west station keeping), momentum dumping, de-orbit (at end of life) and for augmenting orbit transfer. The XIPS operates in two modes, 2.2kW for attitude control and 4.4kW for orbit transfer. There are currently forty 25cm ion thrusters in orbit on 10 Boeing 702 satellites. These thrusters have accumulated an estimated 33,000 hours of successful in-space operation. The normal operation temperature ranges from -50°C to +193°C. Typical 25cm ion thruster performances are summarized in table 4.

Table4. 25cm Thruster Performance

Parameter Low Power High Power

Total Input Power (W) 2067 4215 Thrust (mN) 79 165

Specific Impulse (s) 3400 3500 Electrical Efficiency (%) 87 87 Mass Utilization Eff. (%) 82 84

Beam Voltage (V) 1215 1215 Beam Current (A) 1.43 3.01

The XIPS power controller (XPC) serves as the control and interface system for the 25 cm xenon thruster. The

XPC provides conditioned voltages from the spacecraft bus to the thruster, performs timing and sequencing functions for the thruster start-up and fault protection procedures. Additionally it collects thruster and internal XPC telemetry measurements for ground monitoring. Each XPC is connected to a diagonal pair consisting of one north and one south thruster. While both XPCs are used in nominal operation, one XPC and diagonal thruster pair are capable of performing all maneuvers for the required spacecraft life.

The XPC is powered from the 100-Vdc spacecraft bus, and consists of seven separate power supplies. One

screen, one accelerator, one discharge, two cathode keeper and two cathode heater supplies control all thruster functions. Each power supply has an individual enable/disable control for purposes of sequence and fault control. A separate commandable housekeeping supply provides power to all low power logic, telemetry, and control circuits within the XPC. The XPC is capable of operating the thruster at two different levels of beam current. Table 4 summarizes typical XPC performances.

Table5. 25cm XPC Performance


Total Input Power (KW) 2.2 and 4.5 Bus Input Voltage (V) 100 XPC Efficiency (%) 91% low power

93% high power Size (cm) 20.6x54.1x35.3 Mass (kg) 21.3

30cm Xenon Ion Thruster and Power Processor The 30cm xenon ion thrusters and the associated power processor units (PPU) and digital control interface units

(DCIU) were developed and produced under a contract from NASA GRC for the flight NSTAR system. The thruster operation temperature ranges from -109°C to +158°C. The thruster was designed to operate over a large range of thrust levels as required for the intended deep-space mission applications. The performances at highest and lowest power levels are summarized in table 6.

Table 6. 30cm NSTAR Thruster Performance

Parameter Low Power High Power Total Input Power (W) 490 2310

Thrust (mN) 19 92 Specific Impulse (s) 1950 3280

Electrical Efficiency (%) 82 89 Total Efficiency (%) 38 64

Two NSTAR flight sets were delivered by ETI in December 1997, each consisting of one thruster, one PPU and

one DCIU. The NSTAR EP system was the primary propulsion on the Deep Space One spacecraft that was launched in October 1998. During the mission, the NSTAR EP system achieved more than 16,000 hours of flawless operation. The flight spare NSTAR thruster also operated for more than 30,000 hours and processed 230kg of xenon in a life test at JPL.

In 2005, ETI delivered a flight set of three 30 cm ion thrusters and two PPUs for the Dawn Mission. The

thrusters will be required to operate for up to 21,000 hours each and process up to 160kg of xenon. The Dawn spacecraft is scheduled for launch in 2006.

III. Thruster Manufacturing and Testing Figure 1 shows the major steps of how a thruster is assembled and tested.

Figure1. Manufacturing and test flow.

StructureSub-

Assembly

Harness Sub-

Assembly

Thruster Assembly

Electric & Hi-Pot

Test

Burn in Phase

InitialT/Vac Test

VibePost Vibe Check

Post-Vib T/VacTest

FinalFunctional

Test

Sell-off ReviewShipping

Hollow Cathode Sub-

assemblies

Grid Electrode Sub- Assembly

An exploded view of the final thruster assembly is shown in figure 2.The thruster is comprised of five major sub-

assemblies – discharge cathode, neutralizer cathode, grid electrode ion optics, structural subassembly and harness. Figures 3 to 7 show these 5 subassemblies. The grid electrode ion optics subassembly consists of a screen grid operating at the discharge cathode potential, an accelerator grid at -300 V and a decelerator grid at -20 V. The structural subassembly is comprised of a discharge chamber and a mounting ring both operated at voltages up to 1215 V. The thruster body is enclosed in the ground screens which are at spacecraft ground. Some highlights of the equipment used for assembling the thruster are described in the following sections.



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Figure2. Thruster Assembly Explode View.

Figure 3. Discharge Cathode Subassembly. Figure 4. Neutralizer Cathode Subassembly.



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Figure 5. Grid Electrode Subassembly. Figure 6. Harness Subassemblies.



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Figure 7. Thruster Structure Subassembly. Thruster Manufacturing

Thruster assembly is a highly regimented process designed to ensure quality and traceability. All ion thruster

hardware fabrication and assembly are therefore performed in a dedicated 100K clean room. A variety of special equipment and skills are needed to meet the required process control. For example, figure 8 shows a three axis toolmaker’s microscope equipped with a video monitor that is used for documenting grid electrodes alignments – spacing and concentricity. Magnification power up to 500x and digital readouts allow technicians to verify positioning to within 0.5 micrometers in all three axes. Figure 9 shows a real time x-ray machine that is used to check the integrity of the high voltage ceramic isolators of grid electrodes.



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Figure 8. Three axes Z-Scope. Figure 9. Real Time X-Ray machin. Many of the components used in the thrusters are permanently joined by various welding and brazing techniques. For example, gas tungsten arc welding (figure 10) and orbital welding (figure 11) are used for joining feed-lines and high voltage isolators. Argon gas shielding (in a form of gas stream or bell jar) is used to protect parts from oxidation and contamination during the welding process. A laser welder (figure 12) is used for joining small precision parts when weld penetration and size require precise controls. A resistance welder (figure 13) is used to join iron-based metals where leak tight seals are not required. A brazing process for joining metallized ceramics and for joining dissimilar metals takes place in a vacuum furnace (figure 14).

Figure 10. Gas Tungsten Arc Welder. Figure 11. Orbital Welder. Figure 12. Laser Welder. Figure 13. Resistance Welder.

Figure 14. Brazing Vacuum Furnace.

A variety of in-process tests and analyses are performed during the ion thruster manufacturing to ensure the successful completion of a flight ion thruster. Two stainless steel vacuum chambers with turbo-molecular pumps capable of achieving an ultimate pressure of 10-8 Torr are used for testing up to eight cathode heaters at a time. Heater temperatures are measured with optical pyrometers while the heaters are cycled through acceptance testing. Figure 15 shows 4 heaters during temperature cycling testing. Figure 16 shows 10 individual test chambers pumped with dedicated ion pumps that are used for on-going cathode heater qualification and life testing.



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Figure 15. Cathode Heater Acceptance Testing. Figure 16. Cathode Heater Life Testing. Figures 17 to 19 show the Auger Spectrometer, Scanning Electron Microscope (SEM) with Energy Dispersive X-ray analysis (EDX) and Electron Spectroscopy for Chemical Analysis (ESCA) instruments that are used for contamination analysis, process verification and cathode surface study.

Figure17. Auger Spectrometer. Figure 18. ESCA.

Figure19. SEM – EDX. Ion Thruster Test Facilities and Diagnostics ETI has four 2.7m diameter, 4.6 m long vacuum test chambers built for flight ion thruster acceptance testing.

Each chamber is equipped with ten 1.2-m liquid helium cryopumps that have a total of 300,000 standard liters/s xenon pumping speed. We also have a 6.1 m diameter, 12.2 m long vacuum chamber (Figure 20) with 30 cryopumps which have a total of 1,000,000 standard liters/s xenon pumping speed. This chamber is dedicated for the ion thruster life testing. All of the chambers are capable of producing ultimate vacuums of 10-8 Torr. The thruster under testing is mounted within a liquid nitrogen cooled thermal shroud equipped with resistive heaters, permitting simulation of on-orbit thermal environments.

Figure20. Ion Thruster Life Testing.

Each of the chambers has a water cooled graphite target (Figure 21) in which are embedded 197 Faraday probes allowing characterization of the far field thruster ion current density variation (plume). The system of 197 probes is



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referred to as a thrust vector probe (TVP) because the centroid of the ion current density variation can be used to determine the direction of the thrust vector.

One of the 2.7 m diameter vacuum chambers is equipped with an ExB probe (Figure 22). It measures axial and

radial distributions of singly and doubly charged xenon ions. A Faraday probe (Figure 22) is mounted on the back of the ExB probe head which measures the xenon ion beam profile.



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Figure 21. Photograph of in-vacuum actuator used to position Faraday and ExB probes. Note beam target in background that is instrumented with 197 fixed current density probes that are used to determine the thrust vector direction and thrust loss factor due to ion beam divergence.

ExB PROBE

ExB PROBE

LATERAL MOTION

FARADAY PROBE

THRUSTER VECTOR PROBE

ROTARY MOTION

FARADAY PROBE

Figure22. Faraday and ExB probes.

The photograph in Fig. 21 shows an in-vacuum motion system that is used to position Faraday and ExB probes across the diameter of an ion thruster under test. In addition to linear motion, the motion apparatus is capable of rotating the Faraday and ExB probes into a position where they could be used to characterize the near-field region of an ion thruster beam. The thrust vector probe structure is shown ~3-m behind the motion system.

The Faraday probe shown in Fig. 22 consists of a graphite case that is machined with a sharp-edged orifice with a 0.079 cm2 open area. The graphite case serves as a Faraday shield for a larger diameter graphite disk that is located directly behind the orifice of the Faraday case. The orifice serves as a defining aperture through which high energy beam ions flow before being collected on the disk. The potential of the Faraday case is typically held at -20 V relative to the test facility ground, while the graphite disk is held at -12 V. This biasing scheme is used to repel beam plasma electrons from the disk, while simultaneously preventing Auger (or secondary) electrons from leaving its surface. A current density profile is measured with the Faraday probe by first orienting the probe orifice to be anti-parallel to the ion thruster axis, and then moving it across the face of the thruster while recording the current flowing to the graphite disk as a function of position. The Faraday probe is positioned downstream axially from the center aperture of the thruster by 2.5 cm. The current density is calculated by dividing the probe current by the area of the orifice. A typical Faraday (or current density trace) is shown in Fig. 23 that was acquired with a laboratory model NASA NSTAR thruster fabricated by JPL. The thruster was operated at a beam current of 1740 mA and a net accelerating voltage of 1095 V. An integration of the beam current density trace yielded excellent agreement with measured data (1.1% lower than the measured current or 1721 mA). A flatness parameter (F = < j b > / j b max ) was also calculated for the current density data shown in Fig. 23 and found to be 0.364, which is in good agreement with previous measurement.

0

2

4

6

-20 -10 0 10 20Lateral Position (cm)

Cur

rent

Den

sity

(mA

/cm

2 )

Jb = 1721 mAF = 0.364

Figure 23. Faraday probe trace measured with an NSTAR prototype thruster built by JPL.

A miniaturized ExB probe is also integrated into the turret head of the diagnostics system. It is equipped

with a 5-cm long entrance collimator that has two small orifices at each end. A typical ExB probe trace is shown in Fig. 24. These data were obtained by sweeping the plate voltage to vary the E-field within the ExB filter stage and discriminate between ions of different charge state while monitoring the ion current flowing to the collector electrode located at the back end of the probe. The ExB probe is typically positioned downstream from the ion thruster by 25 cm to 40 cm. The diameters of the holes of the collimator are selected to limit the view of the probe to a region that only contains several (~3 to 6) accelerator grid apertures. This allows one to obtain distributions of the ion charge state profile as the probe is moved across the face of the thruster. The turret head can also be used to rotate the ExB probe to look in different directions relative to the local normal of the ion optics region under inspection, and these data can be used to determine the ion beam divergence characteristics across the face of the thruster. Integrations of the radial variations in charge state and divergence can be used to determine performance parameters that affect thrust and specific impulse calculations.



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Figure 25a contains a plot of typical data collected with the 197 far-field thrust vector probes when an NASA NSTAR prototype thruster was operated at 2.3 KW. It is a plot of the product of ion current density and distance squared to the probe as a function of polar angle (i.e., the angle between the thruster axis and the line between the thruster and a probe). The curve fit shown in Fig. 4a is of the form suggested by Reynolds3 which has been described as a highly collimated, cosine-like distribution with exponential wings. Reynolds curve fits are useful for satellite designers who want to determine the sputter erosion that can occur on a satellite surface that is placed within an ion thruster flow field. The x_bar and y_bar values listed in Fig. 25a correspond to the centroid position measured relative to the center of the thruster vector probe structure. The TVP data can also be integrated to determine the beam current. Typical agreement between the integrated and measured beam current is observed to be <3% once effects of probe orifice plate thickness and charge exchange losses are included. Figure 4b contains a plot that shows how much beam current is contained within a given divergence angle. For the curve fit shown in Fig. 4a, it was observed that a divergence half-angle of 21.5° contained 95% of the total ion beam current. TVP data are also useful for determination of thrust loss factor due to divergence, which is useful for correction calculations of thrust and specific impulse values calculated from the measured beam current and net accelerating voltage.

0

0.2

0.4

0.6

8 12 16 20

Plate Voltage (V)

Col

lect

or C

urre

nt (n

A)

Xe +

Xe ++

VB = 1000 Vj++/j+ = 3.8%

Figure 24. Typical ExB probe trace. Note flat-top features of charge state data that allow simplified calculation of double-to-single ratios and fast radial and angular sampling of beam profile.



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0.0

0.2

0.4

0.6

0.8

1.0

1.2

0 5 10 15 20 25 30

Polar Angle [α ] (o)Fr

actio

n of

Ion

Beam 95% at 21.5o

0.01

0.1

1

10

100

0 5 10 15 20 25 30

Polar Angle [α ] (o)

J*r2 *(

A p c

f cos

( α) )

-1 (A

)

x_bar = -10 cmy_bar = 12.5 cm

JB = 1740 mA (Measured)JB = 1774 mA (Integrated)95% at 21.5o

Figure 25a TVP data for NSTAR prototype. Figure 25b. Fraction of ion beam current contained within a given divergence angle.

We have nine computer controlled thruster power supply consoles that can be operated either in a fully automatic or manual (user controlled) mode. Instrument control and data acquisition are accomplished through National Instruments LabVIEW (Laboratory Viritual Instrument Engineering Workbench), a graphical programming language designed to present a user interface that imitates an actual instrument. Technicians can control thruster power inputs directly at the computer screen using virtual buttons and dials, as well as receive telemetry on all thruster currents and voltages in real time. The console provides a data file for every thruster run. A typical file contains a time averaged set of thruster input and output parameters for every minute of thruster on time. Thruster performance history throughout the whole test profile can be reviewed later. Safety limits are programmed in the computer - the console will shut down the thruster under test if a thruster parameter exceeds its pre-set limit to avoid an operating condition that may be hazardous to the thruster.

Ion Thruster Acceptance Test and Life Test Each flight thruster must pass the acceptance test before the delivery. Testing includes a burn in run, 5 thermal

vacuum cycles, a pre-vibe performance test, vibration, a post-vibe performance test, 3 more thermal vacuum cycles, a final performance test and physical examination. Figure 1 summarizes the test flow. During the acceptance testing, detailed performance data of the thruster are obtained every minute during each run. The thruster’s output operation parameters – beam current, discharge voltage, discharge current, accelerator grid current, decelerator grid current, coupling voltage, discharge and neutralizer cathode keeper voltages along with the calculated operation parameters such as thrust, specific impulse, electrical efficiency, propellant utilization efficiency and thruster efficiency are recorded and thoroughly assessed. Besides these operation parameters, the thruster perveance, backstreaming, thruster operation point sensitivity, thruster vector angle and beam divergence are measured during each performance test.

The typical 13 cm thruster acceptance test run time is about 200 hours and the 25 cm thruster acceptance test is

about 125 hours. Each thruster is random vibration tested in 3 axes with the vibration levels ranged from 3.5 to 8 Grms depending on customer requirements. The vibration frequency range is 20 to 2000 Hz. The test duration is 1 or 3 minutes per axis again depending on the requirements.

The 29th International Electric Propulsion Conference, Princeton University, October 31 – November 4, 2005

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13 cm and 25 cm thrusters were cyclic life tested with a time span that simulated two 15 year mission lifetime with 15% margins. The life tested thrusters were first tested to the acceptance test procedure, followed by more extreme thermal/vacuum and vibration testing and then cyclic life tested. The 13 cm thruster was tested on a 5 hour on and 1 hour off cycle. The 25 cm thruster was tested on a 50 minute on and 30 minute off cycle. Table 7 summarizes the life test status. Some highlights are described in the following sections.

Table7. Summary of Ion Thruster Life Test Data

3 , 2 7 51 6 , 1 4 60 . 5 K W1 3 c m ( Q 1 )

1 2 , 8 4 11 2 , 6 5 82 K W2 5 c m

3 2 42 , 6 8 04 K W2 5 c m

3 , 3 6 92 1 , 0 5 80 . 5 K W1 3 c m ( Q 2 )

C y c l eO p e r a t i o n H o u rP o w e r M o d eT h r u s t e r

3 , 2 7 51 6 , 1 4 60 . 5 K W1 3 c m ( Q 1 )

1 2 , 8 4 11 2 , 6 5 82 K W2 5 c m

3 2 42 , 6 8 04 K W2 5 c m

3 , 3 6 92 1 , 0 5 80 . 5 K W1 3 c m ( Q 2 )

C y c l eO p e r a t i o n H o u rP o w e r M o d eT h r u s t e r

Two 13 cm qualification-model thrusters, Q1 and Q2, were subjected to extensive life testing after completing

their qualification test programs. The tests were conducted in a cyclic manner, with a nominal 5hr ON time selected as being typical of the duration of on-orbit burns. The goal of the cyclic life tests was to achieve 21,000 hrs in 4,200 ON/OFF cycles. In order to complete the tests within a reasonable period of time, they were accelerated by selecting an OFF time of only 1hr between successive ON cycles. Detailed performance data of the Q1 and Q2 thruster were obtained throughout the life test. On average, both thrust and ISP, which are calculated from measured beam voltage, beam current , and total flow rate, degraded 1.2% per 1000 hour for the Q1 thruster; and 0.5 % per 1000 hour, for the Q2 thruster. The 13 cm Q1 life test was terminated at 16,146 hours of operation after the thruster failed to start due to excessive recycles caused by the onset of electron backstreaming. Physical sputtering of the ion extraction electrodes was believed to be the primary life-limiting mechanism at the outset of the life testing. This was evident from the results of the destructive physical analysis (DPA) of the Q1 thruster. The DPA revealed multiple sections of the accelerator and decelerator grids that were completely eroded away. We believe the cause was due to the initial misalignment of screen-to-accelerator grids resulting in sputter erosion of the accelerator grid by the beam ions. This resulted the grid material molybdenum (Mo) being sputtered and deposited onto the screen grid. The sputtered Mo film eventually flaked off as a curved sliver which deflected some of the xenon ions directly into the accelerator grid and milled through the accelerator and decelerator grids. The 13 cm Q2 life test was terminated voluntarily after the thruster completed operation at the 21,058 hours - meeting the life test goal of 21,000 hours of operation. The 25 cm life test began in April of 2000. The life test had two phases – high power and low power. The high power phase testing was modeled after the worst case Boeing 702 spacecraft orbit raising scenario ( i.e. – assume one thruster/XPC pair fails immediately after launch). The high power testing had a duty cycle of 23 hours on, 1 hour off. The testing progressed until the required 2300 hours of high power operation were met (actual testing was completed at 2680 hours) at which time the grid set was removed from the thruster and closely examined. No evidence of significant grid wear was detectable using high power magnification and weight change techniques. In April of 2001 operation switched to low power station keeping operation with a duty cycle of 50 minutes on and 30 minutes off. After every 500 hours of operation, the thruster was subjected a -40°C to +193°C thermal cycle and a complete performance testing. At this time over 13,000 hours of low power operation have been completed.



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Since 2/15/02, (approximately thruster operation hour 2580), the life test has been run by an engineering model XPC and an Inline Data System (ILDS). Previously the testing had been performed with a custom designed test console. Note that the test console still is necessary to conduct functional performance tests or troubleshooting. Computer-averaged data are automatically recorded. The operating parameters recorded are: on-time, chamber pressure,

recycle count, discharge current and voltage, discharge keeper current and voltage, beam current and voltage, accelerator grid current and voltage, decelerator grid current, neutralizer keeper current and voltage, coupling voltage, discharge heater current, discharge heater voltage, neutralizer heater current, neutralizer heater voltage and xenon inlet pressure. In addition, the following performance parameters are automatically calculated: flow, Thrust, specific impulse, Power, electrical efficiency, propellant utilization efficiency and overall thruster efficiency. As of today, no measurable performance parameters changes have been observed - as seen in Figure 26, the thrust ranged from just below 80 to about 80.9 mN meets the thrust requirement of 79.0 mN minimum. Specific impulse (Isp) is above the specification of 3400 s. At 2237 low power run hour, the xenon flow was adjusted to increase discharge voltage to its maximum allowed value. This highest discharge voltage condition was done to maximize screen grid erosion thus providing a conservative lifetime. The Isp increased from above 3400 s to 3600 s when flow was readjusted (figure 27). In summary, after 2680 hours at high power and 13,000 hours of low power operation, the 25 cm thruster remains operational and all performance parameters are in specification. As of today the life testing is continuing.

LIFE TEST LOW POWER PLOTThrust vs. On-Time

20.0

30.0

40.0

50.0

60.0

70.0

80.0

90.0

100.0

0 2000 4000 6000 8000 10000 12000

On-Time (hr)

Thru

st

(mN

)

Figure 26. 25 cm Thrust Life Trend. Note No measurable changes with time

LIFE TEST Q2 LOW POWER PLOTIsp vs. On-Time

2000

2200

2400

2600

2800

3000

3200

3400

3600

3800

4000

0 2000 4000 6000 8000 10000 12000

On-Time (hr)

Isp

(se

c)

Figure 27. 25 cm Isp Life Trend. Note gas flow was adjusted at 2237 hr to life test screen grid erosion at maxima



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IV. Conclusion ETI is a new company with a long history in the technology development of Electric Propulsion. Decades of

research, development, testing and in-orbit experience have led ETI to a comprehensive understanding of the technology and its manufacturing processes. Early on orbit anomalies or failures have been retired through extensive studies and investigations. ETI is committed to growing the application and capability of electric propulsion. In support of this commitment, ETI will continue to pursue opportunities in the development of electric propulsion technology while continuously improving our design and test capabilities. We will also continue to enhance our production processes and designs for our high performance, long life flight thrusters and PPUs for communications satellites and deep space missions.

V. Acknowledgments The accomplishments and progress made in the ion thruster EP technology are the result of the combined talents

of every member of the ETI - EP product group. The authors would like to acknowledge all of them. We also would like to acknowledge Professor John Williams of Colorado State University for his contribution to this manuscript of the ExB and Faraday probes write up, Larry Martin and Andrew Hanna for providing many photos and Brain Buckley for helping to review the manuscript.

References 1 Christensen, J. A. “Boeing EDD Electric Propulsion Programs Overview,” AIAA Paper 2004-3967, July 2004 2 Goebel, D. M., Martinez-Lavin, M., Bond, T. A., and King, A. M., “Performance of XIPS Electric Propulsion in On-orbit

Station Keeping of the Boeing 702 Spacecraft,” AIAA Paper 2002-4348, July 2002 3 T. Reynolds, “Mathematical Representation of Current Density Profiles from Ion Thrusters,” AIAA Paper 71-693, June

1971.

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