45th AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit
st
2 - 545
August
2009, Denver, Colorado Joint Propulsion Conference
AIAA/ASME/SAE/ASEE
AIAA 2009-4815
& Exhibit
AIAA 2009-4815
2-5 August 2009, Denver, CO
Characterization of the NASA NEXT Thruster
M.W. Crofton,* J.E. Pollard,† E.J. Beiting,‡ R. Spektor,§ K.D. Diamant,** X.L. Eapen†† and R.B. Cohen‡‡
The Aerospace Corporation, El Segundo, CA, 90245
and
M.J. Patterson§§
NASA Glenn Research Center, Cleveland, OH, 44135
The NASA NEXT thruster is engineered to be extremely flexible in terms of input power
and specific impulse, while maintaining acceptable efficiency, and embodies a number of
technological advances over previous ion engine systems.
Some of its operating
characteristics are being examined in detail in a study conducted in a test facility at The
Aerospace Corporation. The work in progress is performed under the umbrella of a Space
Act Agreement, with NASA participation in readying thruster hardware for the
measurements. A series of measurements is in progress to verify basic characteristics of
operation and expand on the available database of information. Among the work elements
planned and/or in progress, are in-depth EMI/EMC, plume particle and plasma probe, and
laser diagnostic measurements.
I.
EM
=
EMI/EMC=
HiVHAC =
I
=
Ib
=
Isp
=
MT
=
MT-TOT
=
NEXT
=
NSTAR =
Pin-PPU
=
PM
=
PPU
=
sccm
=
TL
=
TRL
=
Vbps
=
XIPS
=
η
=
Nomenclature
Engineering Model
Electromagnetic Interference/Electromagnetic Compatibility
High Voltage Hall Accelerator
total impulse, Newton-seconds
ion beam current, i.e. current of expelled ions from thruster, Ampere
thruster specific impulse, seconds
total propellant mass flow rate into the thruster system
total propellant mass throughput, kg
NASA’s Evolutionary Xenon Thruster
NASA Solar Electric Propulsion Technology Application Readiness Program
power entering the PPU, kW
Prototype Model (proto-flight hardware)
Power Processing Unit (provides power to the propulsion system)
standard cubic centimeters per minute (of xenon flow)
thrust level, a standard operating point of NEXT
Technology Readiness Level, a maturity scale for evolving technology
beam power supply voltage
Xenon Ion Propulsion System
efficiency
*
Senior Scientist, Propulsion Science Department, M2-341, AIAA Senior Member.
Senior Scientist, Propulsion Science Department, M2-341, AIAA Member.
‡
Senior Scientist, Propulsion Science Department, M2-341, AIAA Associate Fellow.
§
Senior Member of the Technical Staff, Propulsion Science Department, M2-341, AIAA Member.
**
Research Scientist, Propulsion Science Department, M2-341, AIAA Senior Member.
††
Senior Research Associate, Propulsion Science Department, M2-341.
‡‡
Principal Scientist, Space Materials Laboratory, M2-247, AIAA Member.
§§
Senior Technologist for In-Space Propulsion, AIAA Member.
†
Copyright © 2009 by The Aerospace Corporation. Published by the American Institute of Aeronautics and Astronautics, Inc., with permission.
Copyright © 2009 by The Aerospace Corporation. Published by the American Institute of Aeronautics and Astronautics, Inc., with permission.
II. Introduction
T
he NASA NEXT thruster is engineered to be extremely flexible in terms of input power and specific impulse,
while maintaining acceptable efficiency, and embodies a number of technological advances over previous ion
engine systems. Some of its operating characteristics are being examined in detail in a study conducted in a test
facility at The Aerospace Corporation. The work in progress is performed under the umbrella of a Space Act
Agreement, with NASA participation in readying thruster hardware for the measurements.
The non-reimbursable Space Act Agreement (SAA) was drafted by NASA and The Aerospace Corporation to
establish a collaborative measurement program intended to examine the plume particle and field environments of
latest generation NASA ion propulsion technology. While the NEXT ion engine is the current subject of this effort,
subsequent work on Hall thruster technology is anticipated under the same umbrella agreement, with HiVHAC1 the
likely candidate. A series of measurements is currently in progress to verify basic characteristics of NEXT
operation and expand on the available public-domain and internal databases regarding NASA technology and its
potential use on non-NASA spacecraft systems. Among the work elements planned are in-depth EMI/EMC, plume
particle and plasma probe, optical emission and laser diagnostic measurements. This work is of considerable
relevance to future spacecraft integration of the subject thrusters.
NEXT incorporates design improvements not utilized in the previous generation system (NSTAR, the 30-cm
derated ion propulsion system used on Deep Space 1 and Dawn missions). Among these are better magnetic
materials and circuit performance with respect to temperature tolerance and flat ion beam production, more
advanced grid-clear and recycle-recovery systems, better ion optics and cathode technology, higher efficiency and
greater input power flexibility, and higher thermal rejection performance that contributes to lower operating
temperatures and higher power-processing capability.2-4 The ion optics have tighter tolerances with respect to
aperture alignment and gap control, and together with some beam blanking around the periphery they exhibit much
reduced levels of direct grid impingement.2,4 The cathode keeper material is now graphite instead of molybdenum,
extending the life of this component. As a result of numerous design changes working in concert, NEXT has very
high estimated figures for lifetime, total impulse and total propellant processing capability4,5, superior to any other
xenon electric thruster system. However, the range of specific impulse over which NEXT operates efficiently has
limited overlap with competitors of the Hall-thruster type, and due to its relatively high specific impulse normally
achieves lower figures for thrust and thrust-to-power ratio for the same power input. Larger physical size and
weight are also drawbacks for NEXT.
Charge exchange ion erosion of the accelerator grid is expected to be the first thruster failure mode across the
throttling range, with an estimated minimum grid lifetime of 36,000 hours.4,5 Except at the extreme ends of the
basic throttle table, throughput capability has been estimated at about 800 kg of xenon with total impulse of
approximately 3 107 N-s.6 Despite the large throttling range and greater power processing capability of NEXT,
these critical figures are substantially higher than NSTAR performance.7 NEXT capabilities are also improved
relative to the XIPS-25 system8, a second-generation ion engine that benefits from extensive flight heritage and
manufacturing experience, by a lesser margin. For example, XIPS-25 efficiency is between that of NEXT and
NSTAR, while estimated xenon throughputs for NSTAR and XIPS-25 are 20-25% of the NEXT value.9
The basic 40-level throttle table6 (see Appendix) was developed for NASA missions outside of Earth orbit.10
Electrical power available to operate the thruster system would not be constant, due to the varying sun-spacecraft
distance. As a result, the system would need to operate efficiently over a very wide range of input power
parameters, with appropriate output power and specific impulse. Work has recently been performed on an expanded
performance envelope11, indicating that substantial increases in thrust, power processed, and thrust to power ratio
are possible. The precise limits of these parameters have not yet been determined, but clearly the thrust to power
gap between ion engine and Hall thrusters can be narrowed. Particularly when the expanded performance envelope
is considered, the applicability of NEXT is not confined to deep space. Its performance is generally very attractive
for high delta-V earth-orbit missions, due to a large reduction of on-board propellant mass enabled by the high
specific impulse and efficiency.
The NEXT development program at NASA is currently directed toward achieving TRL 6 by the end of calendar
2009, with prototype model thruster and gimbal, engineering model power processing unit and propellant
management assemblies, and control unit simulator comprising the highest planned hardware fidelity until a space
mission is identified. All of this hardware is already built, and most of the requisite testing has been performed 2,4,1215
, therefore the system is presently very close to TRL 6.
An area in which further NEXT work has been needed is that of precise plume particle and field characterization.
Diagnostic instrumentation was integrated with the long duration test16 (LDT) and the multi-thruster array test17
(MTAT). For the LDT, which uses an upgraded EM3 thruster, a Wien filter measures the charge state ratio 4,18 and
2
American Institute of Aeronautics and Astronautics
three Faraday cups measure radial beam current density profiles,4 while six CCD cameras monitor erosion rates and
patterns throughout the test. For MTAT, the plasma density, electron temperature and plasma potential were
mapped in the near field of the thruster array.17
EMI characterization, detailed optical emission and plume contamination measurements, and precise
determination of thrust losses have not previously been performed on NEXT, but are being addressed by the current
project. In addition, many of the NASA measurements have not been performed with the highest fidelity hardware
(Prototype Model thruster and Engineering Model Power Processing Unit).
The overall measurement program at The Aerospace Corporation includes work elements to measure ion beam
flux and divergence, charge state ratios, charge exchange ion flux, plume optical emission spectrum and absolute
flux, radio frequency and microwave absolute emission spectrum plus time-domain emissions, carrier wave
attenuation and phase effects, plume erosion and molybdenum contamination effects, absolute thrust and thrust
correction factors. This report represents work in progress, therefore results of the various measurement programs
will be described in greater detail in the future. Also, additional NEXT measurements are planned that are not yet in
the implementation phase. Current work is being performed with an engineering model NEXT thruster (unit #4,
designated EM4) manufactured by NASA GRC. It has a form and function very close to that of the NEXT
prototype model (PM) thruster manufactured by Aerojet under contract. It is anticipated that a PM thruster will
become available for this project, and critical measurements can be repeated with that system. This PM thruster will
have form, fit, and function identical to a flight article.
III.
Test Apparatus
The vacuum chamber is 2.4 meters in diameter and 10 meters in length. The test facility has eight installed
cryopumps. Four of these are re-entrant and the remaining four are conventional 48-inch-diameter CVI cryotubs.
Figure 1. Photograph of NEXT, installed in The Aerospace Corporation’s facility.
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American Institute of Aeronautics and Astronautics
The re-entrant pumps currently perform more than 25% below their rated speed, probably due to carbon
contamination effects. They are among the oldest pumps on the chamber; two of these have been used in day-to-day
operations since 1993. Each pump is independently operable, so that not all 8 pumps are necessarily operated unless
flow conditions and the desired background pressure demand it. For NEXT operating points of relatively high flow
rate, approximately 150 kl/s can be obtained with all pumps operating.
The NEXT thruster was installed in the facility (see Fig. 1), with core ions directed toward a composite carbon
beam stop mounted on the downstream end dome. More divergent ions impinge on graphoil sheets lining the
cylindrical wall of the chamber. The thruster installation was designed with the flexibility to be positioned either
inside a fiberglass chamber in an anechoic room for EMI work as shown in Fig. 2, or on the axis of the main
chamber as shown in Fig. 1 for all other measurements. NASA personnel managed all aspects of the installation,
which was performed at a high standard with respect to tubing connections, mounting fixtures, flow calibration,
thruster system checkout, etc. NASA has also managed thruster operations during the acquisition of diagnostic data.
For work described in this document, the EM4 thruster was utilized with a NASA-supplied laboratory power
console and xenon feed system. Research grade xenon and rigorous inspection and process control procedures
consistent with those employed for flight hardware were implemented by NASA to ensure xenon purity.
DOWNWARD VIEW
5/2.5
Gnd
RCVR
Pass
Thru
Panel
GATE VALVE
2.4 m diam
x 10 m long
VACUUM
CHAMBER
SEMI-ANECHOIC
ROOM
4.25 m x 5.5 m
PYRAMIDAL
CARBON
BEAM DUMP
Antenna
60 CM
PYRAMIDAL
ABSORBERS
3.4 m
PLUME
THRUSTER
DIELECTRIC
TANK
GRID
FERRITE-BACKED
20 CM
PYRAMIDAL
ABSORBERS
Figure 2. Scale drawing of the electric propulsion electromagnetic compatibility facility.
5/2.5
Gnd
4
American Institute of Aeronautics and Astronautics
Diagnostic equipment in use or planned for future use with NEXT is similar or identical to that employed at The
Aerospace Corporation for other thruster evaluation work (see, for example, Refs. 19 and 20). This equipment
includes a retarding potential analyzer, deflection analyzer, spectrum analyzers and various specialized equipment
for EMI and plume attenuation/phase shift work, laser equipment, a thrust balance under construction, and an optical
spectrometer.
IV.
Plume Characterization
Ion Beam
Plume characterization tests were performed on the NEXT thruster to enable predictions of erosion and
deposition on spacecraft surfaces and to help define stay-out zones for spacecraft appendages. The fluxes of
primary ions and charge-exchange ions were measured for all 40 thrust levels in the basic throttle table as functions
of axial and radial position. The flow angle of primary ions was measured at these locations for a subset of thrust
levels. Instrumented witness plates were used for a preliminary assessment of surface erosion and metallic
deposition in the periphery of the plume. Future measurements will include xenon charge-state distributions vs.
radial position near the exit plane to support modeling of grid erosion and thrust correction factors.
A retarding potential analyzer21 was scanned through the set of axial and radial positions shown in Fig. 3, while
recording the ion current vs. retarding potential at each position. Examples of differential ion flux curves as a
function of the energy per unit charge, E/q, are presented in Fig. 4, indicating the fast primary and slow chargeexchange ions. Because there was minimal ion flux at intermediate energies, a partial scan of retarding potential
from 0 V to 40 V E/q was sufficient to distinguish the fluxes of primary and charge-exchange ions. Partial scans of
1.0E+00
Primary ion flux (mA/cm2)
70
60
50
TL01
TL02
TL03
1.0E-01
TL04
TL05
TL06
40
Axial (cm)
1.0E-02
0
10
20
30
40
50
60
Radial distance (cm)
30
1.0E-01
Primary ion flux (mA/cm2)
20
10
0
-10
0
10
20
30
Radial (cm)
40
50
60
1.0E-02
TL01
1.0E-03
TL02
TL03
1.0E-04
TL04
TL05
1.0E-05
TL06
1.0E-06
1.0E-07
-20
0
20
40
60
80
Axial distance (cm)
Figure 3. Positions and viewing directions for RPA measurements, with examples of primary ion flux vs.
position at thrust levels 01 through 06.
E/q were recorded for all 40 standard thrust levels of the throttle table (see Appendix) to yield spatially dependent
ion flux data, examples of which appear in Fig. 3. Thrust levels 01, 02, and 03 use beam supply voltages of 275,
300, and 400 V, resulting in a much more divergent primary ion beam than the higher thrust levels with beam
voltages of 650 V and above. The charge-exchange flux was about 5 x 10-4 mA/cm2 at all axial positions for thrust
levels 01 through 06.
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American Institute of Aeronautics and Astronautics
The electrostatic deflection analyzer shown in Fig. 5 was used to determine the flow direction of primary ions as
a function of axial position for a subset of thrust levels. At each axial position the repeller potential was scanned
over the primary ion peak while stepping the viewing angle through a range that captured the entire signal. The
mean angle of the flow for thrust level 01 is plotted as data points in Fig. 5 with error bars representing the range
over which the signal was observed. The magenta and yellow curves show what the angle would be if the probe
TL01, axial scan, radial
radial 51
51 cm,
cm, 2/17/09
2/17/09
1.0E-02
cm
(mA/cm2/V)
flux (mA/cm2/V)
RPA differential flux
Primary
57
1.0E-03
51
CEX
44
1.0E-04
38
32
1.0E-05
25
19
1.0E-06
13
6
0
1.0E-07
-6
1.0E-08
0
50
100
150
200
250
300
350
E/q (V)
Figure 4. Differential ion flux vs. E/q measured by the RPA at thrust level 01 for a set of axial distances with
a radial distance of 51 cm.
TL01, radial 51 cm, 2/11/09
100
Vrep
80
Angle (deg)
E/q = 2.67 Vrep
E / E = 2%
FOV = ± 2 deg
Flow
60
Center
Edge
40
Xe+
A
20
0
10
20
30
40
50
60
Axial (cm)
Figure 5. Electrostatic deflection analyzer and measured flow angle vs. axial distance for thrust level 01 with
a radial distance of 51 cm. The direction perpendicular to thruster centerline is defined as 90°.
were viewing the center and the edge of the ion optics. For axial distances of 30 cm or less the flow appeared to
originate at the extreme edge. A similar variation of flow angle with axial position was observed at the other thrust
levels. The results in Fig. 5 determined the programmed viewing angle for axial scans of the retarding potential
analyzer.
Erosion/Deposition
Glass witness plates were exposed to the plume at a subset of the axial positions where the ion flux was
measured. In situ monitoring of erosion and deposition was by conductivity measurements, either from the
progressive removal of a chromium oxide film or from the accumulation of molybdenum produced by sputtering of
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American Institute of Aeronautics and Astronautics
the ion optics. As shown in Fig. 6, thrust levels 01, 02, and 03 produced net erosion at axial distances from 0 cm to
38 cm, consistent with primary ion flux being present at all of these positions. Thrust level 12 was anticipated to
have the highest erosion rate for outer-radius apertures of the ion optics, leading to higher molybdenum levels in the
plume. For TL12 net deposition was detected at axial positions between 0 cm and 25 cm. The presence of
Molybdenum films, radial 51 cm
0.8
60
0.7
50
TL01
40
TL02
TL03
30
3 TL's
20
10
Deposition rate (Ang/hr)
Erosion rate (Ang/hr)
Chromium oxide films, radial 51 cm
70
Preliminary data
0.6
0.5
0.4
TL12
0.3
0.2
0.1
0
0.0
0.0
12.7
25.4
38.1
Axial distance (cm)
0.0
12.7
25.4
Axial distance (cm)
Figure 6. Witness plate measurements of chromium oxide erosion rate vs. axial distance for thrust levels 01,
02, and 03, and of molybdenum deposition rate vs. axial distance for thrust level 12.
molybdenum in the deposited film was confirmed by post-test Auger analysis. The peak deposition rate
corresponded to one monolayer in approximately 4 hours.
The elevated erosion rate at the accelerator grid periphery for TL12 results from poor beam focusing there,
brought about by the combined effects of low beam current density and high total extraction voltage. A partial
masking process used to eliminate most but not all of the beamlets in the annular region between 18 and 20 cm grid
radius was incorporated in the grid set under test. The PM optics and new optics now available for the EM4 reduce
the entire circumference from 40 cm to 36 cm and are expected to produce a significantly lower grid erosion and
plume deposition rate. Grid-to-grid shorting caused by spallation of the accelerator grid and flake formation may
occur at TL12, necessitating an occasional grid-clear operation. Even at TL12, the first failure mode of the ion
optics will be structural failure of the accelerator grid at the center apertures.
Thrust
The NEXT 40-level throttle table contains thrust values calculated from thruster terminal characteristics and
estimated correction factors for off-axis vectoring and multiply charged ions. Insufficient data exist for accurate
calculation of the correction factors, and collection of these data for the full throttle table represents a formidable
experimental task. Direct thrust measurement is anticipated to be a more efficient approach to characterization of the
full throttle table.
Figure 7 is a conceptual illustration of the thrust stand, currently under construction. It is designed as a standard
pendulum with counterweight, and will feature a hardened steel knife edge pivot, a passive magnetic (eddy current)
damper, and a high precision inclinometer for measurement of angular displacement. Gas and electrical connections
will be made as close to the pivot line as possible, with loose or looped leads/tubing to minimize external moments.
Calibration will be performed by applying known moments and measuring the resulting angular displacement.
An elementary analysis of the pendulum, assuming lumped masses and small angular deflections, results in the
following expressions for steady state angular deflection and undamped natural frequency n:
F
g ( mt
mcw
g ( mt
n
[
(1)
l
)
L
mcw
l
L(mcw ( ) 2
L
l
) 1
L ]2
mt )
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American Institute of Aeronautics and Astronautics
(2)
where F is thrust, g is the acceleration due to gravity, mt and mcw are the thruster and counterweight masses, and L
and l are the thruster and counterweight moment arm lengths about the pivot respectively. It is clear from these
expressions that use of a counterweight permits magnification of the angular deflection as well as reduction of the
pendulum natural frequency. The latter may be useful for reducing the pendulum’s response to vacuum chamber
vibration. We intend to target, by adjustment of the counterweight, angular deflections between 1 and 2° over the
complete throttle table (25.5 to 236 mN). In preliminary testing, the temperature controlled inclinometer has
demonstrated repeatability to within 0.002°, and therefore will not preclude achieving our targeted thrust
measurement accuracy of less than 1%.
Beiting has demonstrated the effectiveness of magnetic damping as a non-contact means of reducing pendulum
oscillation.22 A very approximate treatment of the eddy current effect for a single eddy current loop yields the
following expression for the required magnetic field B:
B
L 2R
l
[ (mcw ( ) 2
dH
L
mt )]
1
2
(3)
where d is the effective scale length for the eddy current loop, H is the damper moment arm about the pivot, R is the
resistance to eddy current flow, and is the time constant for the exponential decay of the pendulum oscillation. For
the purpose of estimation, d can reasonably be assumed equal to the size of the magnet pole. Equation 3 provides a
conservative estimate since the aluminum plate (see Fig. 7) will be sized to extend beyond the magnet, and the
magnet (an assembly of four NdFeB permanent magnets) is designed with a field reversal in the path of the plate.
Eddy current loops will be present at the magnet assembly edges, and at the position of field reversal. The magnet
assembly design produces peak field strengths near 0.9 T along the gap centerline, and < 60 seconds is anticipated.
Counterweight
Inclinometer
Pivot
NEXT
N S
Magnet
Aluminum
plate
Figure 7. Thrust stand conceptual diagram.
Electromagnetic Evaluation
Another work element is concerned with electromagnetic studies of the plume. These include radiated emission
spectral scans, radiated emission time domain measurements, and effects on a carrier wave passing through the
plume of the thruster. Carrier wave effects include attenuation, phase shift, and phase noise. At this writing these
measurements are just beginning and only preliminary measurements of radiated emission spectral scan will be
presented.
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American Institute of Aeronautics and Astronautics
Radiated emission from the NEXT thruster is being measured for a subset of six of the 40 standard operating
points. The characteristics of these six points are given in Table 1 (see Appendix also). Because studies of Hall
thrusters have shown that appreciable emission is caused by the neutralizing cathode23, the emission measurements
are being made with the cathode operating in different modes for the two extreme thruster operating modes (TL01
and TL40) and with cathode operating without thruster operation. The cathode is normally operated in “spot” mode
and all results presented here will have the thruster operating in this mode.
Table 1
Mode
0A
0B
0C
1A
1B
1C
2
3
4
5
6A
6B
6C
Thrust
Level
TL00
TL00
TL00
TL01
TL01
TL01
TL05
TL08
TL25
TL37
TL40
TL40
TL40
Cathode
Mode
spot
transition
plume
spot
transition
plume
spot
spot
spot
spot
spot
transition
plume
Vbps
(V)
0
0
0
275
275
275
679
1021
1396
1179
1800
1800
1800
Ib
(A)
0
0
0
1.00
1.00
1.00
1.20
1.20
2.35
3.52
3.52
3.52
3.52
Power
(kW)
0
0
0
0.540
0.540
0.540
1.120
1.520
3.685
4.710
6.860
6.860
6.860
Neutralizer Xe Flow
(sccm)
6.00
3.50
2.00
3.00
2.75
2.50
3.00
3.00
3.50
4.01
4.01
2.50
2.00
The measurements are being taken in The Aerospace Corporation’s Electric Propulsion Electromagnetic
Compatibility Facility (see Fig 2). The facility comprises: a small, all-dielectric tank which is largely transparent to
electromagnetic radiation that houses the thruster and is pumped by the main vacuum chamber; a semi-anechoic
room which surrounds the dielectric chamber and isolates the operating thruster from the ambient electromagnetic
environment and absorbs radiation from the thruster at frequencies higher than 80 MHz; and a set of receivers, low
noise amplifiers, and antennas to detect and record the radiation emanating from the thruster. The receiving
antennas are placed sequentially outside the vacuum 1-m from the thruster centerline. The pressure in the main
chamber during these tests varied between 2.6 x 10-6 and 1.1 x 10-5 Torr (corrected for xenon), depending on the
xenon flow rate from the thruster. The base pressure in the main tank was 9 10-9 Torr. The pressure in the
dielectric tank is about double these values. Details on the design and performance of the EMC facility are
available.24
Table 2
Scan
1
2
3
4
5
6
Band
10 - 150kHz
150kHz - 30MHz
30 MHz - 200 MHz
0.2 - 1 GHz
1 -12.5 GHz
12.5-18 GHz
Antenna
Active Rod
Active Rod
Biconical
Log Periodic
Double Ridge Horn
Standard Gain Horn
6 dB RBW
(kHz)
1
10
100
100
1000
1000
Scan Rate
(sec/MHz)
75
7.5
0.75
0.75
0.03
0.03
Scan Rate
Factor
5
5
5
5
2
2
A composite view of the 36 vertically polarized (E field perpendicular to the axis of the plume) scans of emission
from the NEXT thruster for the six operating conditions given in Table 1 is shown in Figure 8. We emphasize that
these data are preliminary because it is suspected that some scans are contaminated with intermittent backgrounds
emission. This is a log-log plot of the electric field from 10 kHz to 18 GHz measured from a distance of 1 meter. In
this plot a change of 20 dB in the vertical axis is a factor of 10x change in electric field. These data were taken
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American Institute of Aeronautics and Astronautics
under MIL-STD 461E RE102 specifications. Salient scan information is given in Table 2. The frequency scan rate
used was slower (by the factor shown in the last column) than the fastest permitted by 461E to ensure that peak
emissions were recorded. Note from this table the bandwidth increases by 10x at several frequencies (as required by
461E), which explains the discontinuities in amplitude seen in the data in Figure 8 at these frequencies.
In the Figure, the blue curve at the bottom of the graph is the sensitivity of the instrumentation. The lower black
curve is the background emission (thruster off) in the anechoic room taken at an arbitrary time before or after the
scans taken with the thruster operating. The red line is the MIL-STD 461E limit, which is given for reference but is
seldom relevant for a specific spacecraft. The remaining curves are the emission from the thruster for the six
conditions defined in Table 1.
Although the density of data in this figure is too high for detailed analysis, it is useful to obtain an overview of
the emission. First, we note that there is no emission above the background at frequencies between 1-18 GHz. This
is in stark contrast with the emission from Hall thrusters, which show significant radiation in the L-, S-, and Cbands, i.e. 1-8 GHz (see Ref. 23 and references therein). Second, similar to the Hall thrusters, the radiation from the
NEXT thruster is significantly above the 461E limit at frequencies between 10 kHz-20 MHz. We also see that
below a few MHz, the emission scales with power in that the emission for thruster modes TL1, TL5, and TL8 (0.51.5 kW) is significantly below that of modes TL25, TL37, and TL40 (3.7-6.9 kW).
Interpretation of the data between 20 MHz-1000 MHz is problematic due to a possible high intermittent
background, i.e. it may or may not appear in the background scan and can randomly appear in a thruster scan. The
background in this region is high because of the passage of radiation from the main vacuum chamber into the
anechoic room through the 1-m diameter plume exhaust port. The carbon fiber grid used to attenuate this
background leakage24 into the anechoic room by approximately 20 dB was not in place during these measurements.
This background makes the interpretation of the data at these frequencies ambiguous and additional data will be
acquired with the grid in place.
The data in the first scan listed in Table 2 (10 kHz and 150 MHz) do not suffer from this possible background
problem. Figure 9 shows a linear frequency plot of this scan for all six thruster modes studied where the resolution
bandwidth is a constant 1 kHz. This scan registered the highest radiation levels observed for this thruster. This
spectral segment illustrates the detail available in these scans and is interesting because of the strong quasi-periodic
coherent emission appearing near 25, 45, 75, 95, 118, and 142 kHz for all thruster modes studied. This emission is
strongest for the TL37 and TL40 thruster modes for the lowest four frequencies but highest for the three lowest
Figure 8. Log-Log plot of emission from NEXT thruster from 10 kHz to 18 GHz.
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American Institute of Aeronautics and Astronautics
Figure 9. Linear-Log plot of emission from NEXT thruster from 10 kHz to 150 kHz.
power thruster modes for the highest two frequencies. The switching frequency of the beam supply is near 22 kHz
and switching frequency of the discharge, neutralizer and accel power supplies are all near 50 kHz. It is likely that
these emissions are caused by the harmonics and mixing of these frequencies. Switching frequencies are a
characteristic of the commercial power supplies used in the laboratory power console, and the associated noise is not
representative of what EM or flight model PPU hardware will produce. The apparent detection of these noise
sources provides additional impetus for further testing with higher fidelity equipment.
Scan data such as these are useful for a survey measurement of the emission of the thruster. When integrating
this thruster with a particular spacecraft payload, more detail measurements are required, including measurement in
the time domain for the frequencies were the payload has vulnerabilities.
Optical Emission, general considerations
The purpose of the optical spectroscopic measurements is to characterize the absolute plume light emission and
to investigate the feasibility of erosion analysis. Characterization of the absolute emission is needed for potential
light sensitive payloads, such as telescopes and photon detectors. For such measurements we are interested both in
the total light intensity and the spectral nature of the emission. For a typical spacecraft configuration where the
thrusters are placed on the opposite (or adjacent) side from the payload the light sensitive instruments are shielded
from the light emanating from the thruster itself. However, the plume of the thruster may extend far enough to have
a line of sight to the instrument. In general, one would want to determine how much light is produced in each region
of the plume. Such a measurement would allow an accurate determination of the absolute light intensity at an
arbitrary payload location. This approach, however, is unfeasible since spectroscopic measurements produce the
line-integrated intensity.
A simple optical calculation can show that spatially-resolved optical emission measurements require at best a
very fast optical system, and at worst cannot be performed at all. Figure 10a shows a single-lens system, where an
object o, which is located at a distance lo from a thin lens of diameter d and focal length f is imaged onto a photo
detector of diameter i, which is located at a distance li from the lens. Since a plasma plume constitutes a distributed
light source, we can imagine that another region at a closer distance to the lens lo’ will be mapped at some distance
behind the photo detector. Thus, the photo detector will collect less light from that region. The decrease in the
amount of light from that area measured by the detector will roughly be proportional to the detector area (¼πi2)
11
American Institute of Aeronautics and Astronautics
divided by the area of the light cone at the detector position (¼πi’2), as indicated by the diagram. Using a few basic
geometrical optics equations we can derive a formula for the light detection at an arbitrary distance lo’ around the
object o,
i2
i '2
2
l
d
1
1 2 i'
o
lo
li
f
li
lo
.
(4)
Figure 10b shows a sample calculation using Eq. 4 with the lens system employed for the reported measurements,
and shown in Fig. 11. The x axis of Fig. 10b is the normalized distance between the lens (x = 0) and the object of the
system (x = 1). The right axis shows that the amount of the detected light decreases as one moves away from the
object position. The left axis shows that the total light collected by the detector is about 9 times the light collected
from the plasma disk located at the object distance lo. Furthermore, there is significant light collection even at the
FIG. 10: Extended light source calculations. Panel a) shows a schematic of a one-lens optical system. Panel b)
shows a sample calculation of the light detected from an extended source, such as plasma plume, as a function
of distance from the lens using Eq. 4. Here, the x axis is the normalized distance between the lens (x = 0) and
the object (x = 1), while the right axis shows fraction of collected light, and left axis shows total light as a
fraction of light collected from the source at x = 1.
FIG. 11: Experimental setup for measuring NI-UV spectra emitted by the NEXT thruster.
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American Institute of Aeronautics and Astronautics
midpoint between the lens and the object. Thus, with the lens focused in the middle of the thruster beam and light
collection across the beam (with beam diameter of ≈ 20cm) there is significant contribution of light from the entire
line of sight across the beam.
Equation 4 implies that a faster optical system can in principle produce a measurement with sufficient spatial
distribution. However, in reality that expression provides only the best case scenario, because it ignores vignetting.
An analytical expression similar to Eq. 4 that includes vignetting requires challenging computations and is beyond
the scope of this paper. It is quite possible that once that effect is taken into account one will find that the detected
light is homogeneously distributed throughout the line of sight across the plume, and thus a localized measurement
is simply impossible.
An alternative method of measuring absolute emission would be to simulate the light sensitive payload by a
diode photo detector placed and oriented similarly to the simulated payload relative to the thruster and without any
additional optics, such as focusing lenses. By moving such a detector to various locations one can obtain a measure,
albeit imperfect, of the impact on the payload. The spectral nature of the light radiation can be measured by a
collocated spectrograph, or a selection of band-pass filters.
The other purpose of the spectroscopic measurements is to determine the feasibility of quantitative non-invasive
grid erosion estimation. Such measurements have previously been attempted with various success. Rock et al.25 have
developed a Kinetic Steady-State (KSS) model for erosion rate of Molybdenum grids using the MoI 379.8 nm line.
Domonkos and Stevens26 evaluated the KSS model for erosion studies of the NSTAR thruster. Same MoI line has
also been detected by Crofton27, and thus seems to be a reliable feature of the molybdenum gridded thrusters.
Consequently we should expect to see the 379.8 nm line in the emissions of the NEXT thruster as well. Application
of the KSS or other erosion models should allow us to estimate erosion rate of the NEXT thruster grids without
performing the time-consuming life test. Another and more accurate way to estimate the amount of Molybdenum in
the thruster plume is to perform an LIF measurement. Feasibility of such a measurement has been successfully
demonstrated by Crofton et al.28
Figure 12. Spectrum obtained at the thruster exit plane under TL37 conditions, with identification of
prominent lines.
Spectroscopic Measurements
Preliminary spectroscopic investigation of the NEXT thruster emission has been performed during the test
campaign at The Aerospace Corporation. A schematic of the experimental setup is shown in Fig. 11. Measurements
were performed with a single-lens system (f = 100), which was placed outside of the vacuum chamber next to a 4
inch diameter BK7 window with a direct view to the thruster and its plume. The lens focused the light from the
13
American Institute of Aeronautics and Astronautics
centerline of the thruster plume onto an optic fiber, which was fed into an HR4000 spectrograph. The spectrograph
recorded the light emission from 200 to 1000 nm. The other end of the optic fiber was placed on a translation stage
such that the light from various points along the plume could be collected. Light emanating from within the thruster
through the grids (along the centerline) could also be collected by moving the fiber into the appropriate position. In
the following discussion 0 cm corresponds to the position where the outer grid meets the thruster body. At the
centerline the grid extends to ~2.35 cm. Beyond that point the collected light was assumed to come from the
thruster plume. For these preliminary measurements we used a BK-7 vacuum window, and thus most of the light
below 400 nm was not transmitted to the spectrograph, as can be clearly seen in the sample spectrum shown in Fig.
12. The purpose of these preliminary measurements was to survey the spectrum for the planned detailed absolute
emission measurements.
During the test campaign we collected emission at the following thruster operating points: TL01-12, TL19,
TL29, and TL37-38 (see Appendix). Here we will present data only from TL09, TL19, TL29, and TL37. These
thrust levels correspond to the same beam voltage (1179 V) and variable beam currents, increasing from 1.2 to 3.52
A for TL09 to TL37 respectively.
The sample spectrum shown in Fig. 12 was obtained at TL37. The measurement was taken at the thruster exit,
corresponding to axial position of 2.35 cm. The spectrum presents an expected combination of the prominent Xe
neutral and ion lines. Furthermore, some Ti and Mo lines were tentatively identified, however further analysis is
needed to verify their identities. Unfortunately, the 379.8 nm MoI line would not be observable due to the
transmission properties of the window. We anticipate the observation of this line once a quartz window has been
installed. Using the quartz window we plan to observe the functional dependence of the MoI line on the various
thrust levels, and use the KSS model to estimate erosion rates. Furthermore, in addition to the spectroscopic
measurements we hope to investigate spatially-resolved grid erosion with the LIF technique.28
Figure 13 shows spatial variation of xenon neutral (881.9 nm) and ion (529.7 nm) lines. Emission of both lines is
5 to 10 times stronger in the thruster than in the plume. Furthermore, the maximum intensity at all thrust levels,
except at TL09, occurs at the thruster exit (2.35 cm). At a given position the line intensity (for both ion and neutral
lines) increases linearly with the beam current.
With the preliminary spectroscopic measurements completed, we plan to perform an absolute intensity
measurement with diode photodetectors. These detectors will be mounted on a swing arm behind the thruster. The
swing arm will allow placing the detectors at an arbitrary position and orienting with respect to the thruster.
Collocating multiple detectors with various optical band-pass filters will enable us to determine a rough absolute
FIG. 13: Xenon ion and neutral line intensity as a function of beam current for beam voltage V b=1179 V.
Radiation intensity from both species scales linearly with the beam current. Intensity increases toward the
thruster exit and then drops an order of magnitude in the plume.
spectral signature of light emission at the payload location. In addition to that, we plan to place an optic fiber to
collect the spectrum similarly to the already performed preliminary measurements. The spectrograph can also be
calibrated for the absolute intensity measurements.
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American Institute of Aeronautics and Astronautics
V.
Conclusions
A comprehensive plume characterization study is in progress that will significantly improve the knowledge base
concerning particles and fields associated with the NEXT ion engine plume. This is a collaborative effort intended
to result in the evaluation of selected NASA electric thruster technology that is of interest to the space propulsion
community. Preliminary results include new EMI, optical emission, ion beam, contamination, and charge exchange
ion data. Further measurements and more detailed reporting are planned.
VI. Appendix
NEXT throttle table at beginning of life, from Ref. 6b.
TL
Vbps, V
F, mN
Isp, sec
TL40
TL39
TL38
TL37
1800
1567
1396
1179
236
221
208
192
4190
3910
3690
3395
TL36
TL35
TL34
TL33
1800
1567
1396
1179
208
194
184
169
4150
3875
3660
3360
TL32
TL31
TL30
TL29
TL28
1800
1567
1396
1179
1021
181
169
160
147
137
4150
3875
3660
3360
3125
TL27
TL26
TL25
TL24
TL23
1800
1567
1396
1179
1021
158
147
139
128
119
4100
3830
3615
3325
3090
TL22
TL21
TL20
TL19
TL18
1800
1567
1396
1179
1021
134
125
118
108
101
4310
4025
3800
3490
3250
TL17
TL16
TL15
TL14
TL13
1800
1567
1396
1179
1021
107
99.9
94.3
86.7
80.6
4190
3910
3690
3395
3155
TL12
TL11
TL10
TL09
TL08
TL07
TL06
TL05
TL04
TL03
TL02
1800
1567
1396
1179
1021
936
850
679
650
400
300
80.2
74.9
70.7
65.0
60.4
57.8
55.1
49.2
48.1
37.3
31.9
4000
3735
3525
3240
3015
2885
2745
2450
2400
1860
1590
TL01
275
25.6
1410
Pin-PPU,
ηPPU
kW
MT = 5.76 mg/sec; Ib = 3.52 A
0.71
7.240
0.948
0.70
6.385
0.948
0.69
5.780
0.945
0.68
5.000
0.941
MT = 5.12 mg/sec; Ib = 3.10 A
0.70
6.390
0.948
0.69
5.645
0.947
0.68
5.095
0.947
0.67
4.420
0.941
MT = 4.46 mg/sec; Ib = 2.70 A
0.70
5.570
0.950
0.69
4.930
0.948
0.68
4.450
0.947
0.67
3.870
0.941
0.65
3.435
0.938
MT = 3.92 mg/sec; Ib = 2.35 A
0.69
4.870
0.948
0.68
4.310
0.947
0.67
3.900
0.944
0.66
3.380
0.941
0.64
3.010
0.937
MT = 3.16 mg/sec; Ib = 2.00 A
0.71
4.215
0.950
0.70
3.745
0.947
0.69
3.400
0.944
0.67
2.955
0.941
0.65
2.635
0.938
MT = 2.60 mg/sec; Ib = 1.60 A
0.68
3.420
0.950
0.67
3.040
0.947
0.65
2.765
0.944
0.64
2.410
0.941
0.62
2.160
0.935
MT = 2.05 mg/sec; Ib = 1.20 A
0.65
2.575
0.948
0.63
2.285
0.946
0.62
2.075
0.944
0.61
1.810
0.943
0.59
1.625
0.936
0.58
1.520
0.935
0.56
1.420
0.931
0.53
1.215
0.922
0.52
1.185
0.918
0.43
0.890
0.884
0.37
0.770
0.868
MT = 1.85 mg/sec; Ib = 1.00 A
0.33
0.630
0.868
ηt
Lt, khr
MT-TOT,
kg Xe
I, x107
N-s
0.671
0.663
0.653
0.637
37.6
37.6
37.6
40.5
780
780
780
840
3.20
2.99
2.82
2.80
0.663
0.654
0.646
0.629
43.0
43.0
43.0
46.3
795
795
795
855
3.22
3.01
2.84
2.81
0.662
0.653
0.644
0.626
0.610
49.1
49.1
49.1
52.9
65.7
790
790
790
850
1055
3.20
2.99
2.83
2.80
3.23
0.652
0.642
0.632
0.616
0.600
56.5
56.5
56.5
60.8
75.5
795
795
795
860
1065
3.21
2.99
2.83
2.80
3.23
0.670
0.658
0.646
0.628
0.609
66.4
66.4
66.4
71.5
88.9
755
755
755
815
1010
3.20
2.98
2.82
2.79
3.22
0.642
0.630
0.618
0.598
0.577
83.6
83.6
83.6
90.1
112
785
785
785
845
1050
3.22
3.01
2.84
2.81
3.25
0.611
0.600
0.589
0.571
0.550
0.538
0.523
0.486
0.479
0.381
0.323
110
110
110
119
147
196
293
366
212
65.2
47.2
815
815
815
875
1090
1445
2160
2705
1565
480
350
3.18
2.97
2.80
2.78
3.21
4.07
5.80
6.49
3.68
0.88
0.54
0.282
68.0
500
0.63
ηSystem
15
American Institute of Aeronautics and Astronautics
Acknowledgments
This work was supported under The Aerospace Corporation's Independent Research and Development Program
and NASA’s In Space Propulsion Technology Program, Science Mission Directorate. The authors thank Kevin
McCormick of NASA Glenn Research Center for outstanding technical support throughout the process of thruster
system installation, and Carrol Gardner and Kevin Dorman of The Aerospace Corporation for invaluable assistance.
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