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NASA/TM—2004-213194

The High Power Electric Propulsion (HiPEP) Ion Thruster

John E. Foster, Tom Haag, and Michael Patterson

National Aeronautics and Space Administration

Glenn Research Center
Cleveland, Ohio 44135

George J. Williams, Jr.

Ohio Aerospace Institute

Brook Park, Ohio 44142

James S. Sovey

Alpha-Port, Inc.

Cleveland, Ohio 44135

Christian Carpenter

QSS, Inc.

Cleveland, Ohio 44135

Hani Kamhawi, Shane Malone, and Fred Elliot

National Aeronautics and Space Administration

Glenn Research Center
Cleveland, Ohio 44135

Practical implementation of the proposed Jupiter Icy Moon Orbiter (JIMO) mission,

which would require a total delta V of approximately 38 km/s, will require the development
of a high power, high specific impulse propulsion system. Initial analyses show that high
power gridded ion thrusters could satisfy JIMO mission requirements. A NASA GRC-led
team is developing a large area, high specific impulse, nominally 25 kW ion thruster to
satisfy both the performance and the lifetime requirements for this proposed mission. The
design philosophy and development status as well as a thruster performance assessment are
presented.

Introduction

igh power nuclear electric propulsion (NEP) systems are an enabling technology that has the potential to allow
for the intensive exploration of the outer planets. Additionally, ample power for mission science and

communications would be available via the nuclear power source. This on-board power source eliminates the solar
flux penalty that limits the practical reach of solar electric propulsion to the inner solar system. The ample power of
a NEP system would allow for continuous thrusting over the period of the expected life of the reactor, which for
liquid metal cooled systems such as the SP-100 ranges between 7 and 10 years.

The present state of the art in high efficiency, high specific impulse, electric propulsion systems is embodied in

the NASA’s Solar Electric Propulsion Application Readiness (NSTAR) ion thruster. This thruster operated for over
16,000 hours in space and over 30,000 hours during an extended life test (ELT).

While significant in its own regard,

the use of NSTAR technology, is insufficient to satisfy life and performance requirements for long duration missions
to the outer planets. Depending on the mission, required high specific impulse engine thrusting times can be a
significant fraction of the lifetime of the reactor. For example, the proposed Jupiter Icy Moon Orbiter mission has a
delta-V requirement of at least 38 km/s. Ion thruster systems used for such an application will be required to operate
continuously for perhaps as long as 7-14 years.

3-5

Such long continuous operation times place stringent lifetime

requirements on thruster components and subsystems.

NASA/TM—2004-213194

In an effort to develop the engine technology capable of satisfying the demanding performance and lifetime

requirements for NEP class mission, a NASA Research Announcement (NRA) was issued. The NRA called for the
development of a long life, high power, high specific impulse propulsion system that could satisfy requirements for
nuclear-electric propulsion missions to the outer planets and beyond. The initial requirements as defined by the NRA
are listed below:

• Specific impulse = 6000 to 9000 seconds

• Thruster Efficiency >65%

• NEP system specific mass = 30 kg/kW

• Propellant throughput = 100 kg/kW

Thruster requirements have since changed becoming more stringent with power, specific impulse, and

throughput/life so as to satisfy Jupiter Icy Moon Orbiter (JIMO) mission requirements.

This translates into very

long thrusting times (7-14 yrs) depending of specific impulse.

The research announcement has led to the development of two gridded ion thruster systems. The High Power

Electric Propulsion (HiPEP) project led by NASA Glenn Research Center (GRC) and the Nuclear Electric Xenon
Ion Thruster System (NEXIS) project led by NASA’s Jet Propulsion Laboratory (JPL) responded to NASA’s
research announcement to develop such a propulsion system. The NEXIS thruster system approach was to leverage
the heritage embodied in the NSTAR and the NASA Evolutionary Xenon Ion Thruster (NEXT) thruster.

6,7

The

HiPEP approach infuses technologies such as a rectangular discharge chamber, long life grid materials, and flexible
discharge plasma production (DC or microwave electron cyclotron resonance). This approach addresses the primary
thruster failure modes. The development of these technologies was seen as necessary since the life requirements will
require significant hurdles over the state of the art. As such, the development of these technologies is focused on
mitigating both cost and schedule risk.

The focus of this report is to give a general description of the HiPEP thruster, addressing the philosophy of the

approach, the implementation, and analysis of data to date characterizing engine operation. Details regarding HiPEP
project structure and overall scope may be found in the companion reference.

Based on findings from mission analysis for 8000 s specific impulse, a 25 kW design point was chosen as the

baseline operation point. In order to respond to potential changes in the JIMO specific impulse requirement, a 6000 s
specific impulse, 25 kW was selected as a secondary design point.

II.

Design Considerations

In general, thruster lifetime is limited by essentially five potential failure modes: 1.) discharge cathode failure,

2.) neutralizer cathode failure, 3.) electron backstreaming, 4.) erosion induced structural failure of the ion extraction
grids, and 5) formation of an unclearable short between grids. The failure mechanism of the cathode assembly is
multi-faceted in that subcomponents such as heater failure and keeper assembly erosion can ultimately lead to
component failure. Cathode failure modes can be loosely grouped into two general areas: physical erosion via
sputtering and emitter element depletion of low work function material. Physical sputter erosion of the cathode
assembly occurs because the cathode is constantly subjected to ion bombardment from the surrounding discharge
plasma. Emitter failure is related to thermo-chemical processes that render the cathode incapable of supplying
electrons even if other conditions such as thermal environment and pressure are adequate. Provided gas cleanliness
protocols are followed (eliminates emitter poisoning), emitter failure occurs after long operation times because of
the depletion of work function lowering impregnates at the emission sites. Additionally, the high temperature
formation of inert or emitter pore-plugging compounds also effectively reduce the supply of work function lowering
impregnates. Conventional ion thruster hollow cathodes have a demonstrated lifetimes of order 28,000-30,000

2,9

hours. Longer hollow cathode lifetimes need to be demonstrated for these components to be compatible with those
missions requiring continuous thruster operation in excess of that demonstrated to date. Failure mode 3 is related to
electron backstreaming which occurs when accelerator grid apertures widen (at fixed accelerator grid voltage) due to
erosion. When the aperture is sufficiently large, the positive potential associated with the screen grid can “leak”
downstream of the ion optics assembly. When this occurs, electrons from the beam plasma can actually backstream
into the engine achieving energies approximately equal to the beam voltage. This energetic beam of electrons can
quickly overheat or damage the discharge cathode. Failure mode 4 is associated with severe grid erosion. Sputter
erosion of the ion extraction grids can ultimately lead to thruster failure. Erosion of these components occurs
primarily by charge exchange erosion. If the beamlets are not well focused, erosion due to direct impingement can
also occur. Over time, these ion milling processes lead to structural degradation of the ion optics assembly, leading
to poorer discharge performance over time and ultimately the cantilevering of one electrode into the other, giving

NASA/TM—2004-213194

rise to a short and thereby terminating beam extraction. One potential solution to this problem is the use of a
magnetic grid.

Potential design solutions also exist for increasing the lifetime of the ion optics by using different

electrode materials such as titanium

or carbon

12-13

or by simply increasing the electrode thickness.

Failure mode 5

involves unclearable shorts between the grids. The formation of large conducting flakes formed either from the
erosion of the ion optics electrodes or by erosion of the discharge cathode assembly also can lead to ion optics
failure. If the conducting flake were to bridge the gap between the high voltage ion optics grids, the resulting short
would also terminate beam extraction. Another grid shorting mechanism is caused by unattached debris from
spacecraft surfaces shorting the grids. This event has the highest probability of occurring during the launch phase.

As mentioned, the primary failure modes of the thruster are associated failure of the discharge cathode and the

ion optics. As lifetime is the key parameter, the HiPEP approach necessarily focuses on greatly exceeding the state
of the art for the various ion thruster component technologies. The HiPEP thruster design and development effort
focuses on the elimination of failure modes.

Plasma Production

The HiPEP project approaches the main discharge plasma production issue with a two prong approach: DC

hollow cathode and microwave electron cyclotron resonance (ECR) plasma generation. Both DC and microwave
approaches have been developed. The DC plasma generator is the HiPEP thruster baseline approach. The DC
approach utilizes a NEXT-like discharge cathode.

Discharge current requirements for the 8000 s specific impulse

design point are consistent with the operating range of the NEXT thruster discharge cathode assembly. This
discharge cathode emitter is significantly larger than the NSTAR emitter insert. The larger emitter increases device
lifetime by virtue of the fact that the impregnate reservoir is larger. The larger insert and orifice size accommodate
higher emission current densities than the NSTAR design, with emission currents in excess of 40 A. This emission
current range provides significant margin at the 8000 s specific impulse operating point. The discharge cathode
keeper, which serves as a cathode physical shield against sputtering ions, is subject to erosion and can give rise to
potential failure mechanisms such as formation of large, conducting flakes. Complete erosion of the keeper face
plate, as was observed in the NSTAR 30,000 hr extended life test.

In order to minimize cathode assembly erosion

due to physical sputtering, the main discharge cathode will utilize a graphite keeper. Because the sputter threshold of
xenon ions on graphite is greater than the energies of the expected ion flux (discharge voltage~25 V), the use of the
graphite keeper should eliminate this failure mechanism.

The backup plasma generation approach involves the complete replacement of the hollow cathode assembly

with an electrodeless plasma production.
Microwave ECR has been investigated under the
HiPEP project as an approach to eliminate the
potential discharge cathode failure mechanisms.
This technology has been demonstrated as a viable
plasma production option.

16-20

Indeed, ECR ion

thruster technology has been used as the primary
propulsion for the MUSES-C asteroid rendezvous
mission.

19-20

This electrodeless plasma production

approach, depicted in Fig. 1 utilizes microwaves
that heat electrons resonantly in the presence of a
magnetic field. At this resonance, the electrons can
gain energy continuously. This resonant process
takes place on surfaces of constant magnetic field
that are established by the magnetic circuit. The
hot electrons produced during this process ionize
neutral gas, thereby generating the discharge
plasma completely electrodelessly. The process
takes place away from discharge chamber walls,
thereby minimizing wall erosion. Plasma potentials
associated with ECR plasmas are typically
significantly less than conventional hollow cathode
devices.

In this regard, the sputtering of the

upstream surface of the screen grid can be virtually
eliminated as the ions will strike the grid at
energies associated with the Bohm speed.

Figure 1. Conceptual depiction of electron cyclotron
resonance heating. Resonance occurs when the microwave
frequency

is equal to the electron cyclotron frequency,

Magnetic field lines

Microwaves

NASA/TM—2004-213194

The implementation of the microwave ECR approach for HiPEP employs the use of a slotted antenna. The

slotted antenna affords the opportunity for distributed plasma production. Distributed plasma production yields
uniform plasma density profiles at the optics exit plane, resulting in very flat beam profiles. Recall flat beam profiles
are desirable in that they circumvent issues such as a reduced perveance limit and accelerated, localized accelerator
grid wear on centerline. The microwave energy source proposed by HiPEP for the thruster application is a klystron.
It is expected that the klystron should have lifetimes of order that of space qualified tubes such as the traveling wave
tube (TWT). TWTs have demonstrated on-orbit lifetimes in excess of 144 kHrs.

21-23

The HiPEP main discharge chamber is rectangular in geometry and is designed to accommodate either the

baseline hollow cathode plasma production approach or microwave plasma production. The target plasma
production efficiency for the HiPEP engine < 200 W/A while the design discharge chamber propellant utilization
target is > 90%. Both plasma production approaches have been demonstrated with the rectangular discharge
chamber. The thruster, shown in Fig. 2(a), is large, with an ion extraction exit plane measuring 41 x 91 cm. The
large ion extraction area allows the thruster discharge chamber to operate at a lower current density than
contemporary thrusters. Reduced beam current density reduces grid wear rates.

To achieve the target discharge chamber efficiency and propellant utilization goals, magnetic field calculation

software was used to guide in the design and optimization of the ring cusp magnetic field geometry chosen. Ring
cusp magnetic circuit electron containment schemes have been shown to be very efficient.

In the case of the HiPEP

thruster, the shape of the magnetic rings range from circular to hybrid circular-rectangular to rectangular. Such ring
geometries are necessary to accommodate the discharge chamber shape. The discharge chamber itself is made of
non-ferrous steel. High field strength, rare earth magnets comprise the magnetic circuit. The magnetic circuit design
accommodates the large volume plasma
production necessary for thruster operation at
the design points. Regions of low magnetic
field strength, away from the walls, comprise
a significant fraction of the internal discharge
chamber volume. The termination plane of
the discharge chamber is relatively field free
and thus offers ease of flow of plasma ions to
the ion extraction grids.

Discharge Chamber Shape, an Aside

It should be pointed out that the

rectangular shape of the discharge chamber
and associated plasma screen is well shaped
for multiple thruster installations. Many
thrusters can be installed adjacent to each
other, forming a dense cluster of aligned ion
beams. For multi-megawatt spacecraft, close
packing may be desirable in order to
consolidate structural mass, and minimize
spacecraft appendages. Close packing may
also enable collective beam neutralization,
either with redundant neutralizers, or a few
high capacity plasma contactors. Thermal
management issues associated with close
packing of engines can be addressed by
simply mounting the thrusters in a manner
such that the back-plate of the engines is
exposed to the vacuum of space, thereby
allowing the engines to freely radiate out of
this plane. Fig. 2(b) depicts an artist
conception of a possible HiPEP packing
approach for rectangular geometry ion
engines.

From a power processing standpoint, a

primary attribute of the rectangular geometry

(a)

HiPEP Thruster

Neutralizer

(b)

Figure 2. (a) HiPEP rectangular thruster featuring large area
ion optics. (b) Conceptual depiction of HiPEP thruster pod:
Rectangular thruster geometry is particularly amenable to
“pack and stack” integration.

NASA/TM—2004-213194

is scalability. The ability of the thruster discharge chamber to grow in size to accommodate higher power operation
is an important thruster attribute in that it provides the flexibility to accommodate changes as mission requirements
vary. Such flexibility also allows the engine to be applicable to a range of missions requiring different power levels.
The rectangular geometry can accommodate significant increases in cross sectional area by simply increasing its
lateral dimension. Because the internal magnetic circuit is rectangular in geometry as well, all that is required of the
magnetic circuit to be consistent with increasing of the lateral dimension is simple lateral “stretching.” In this
respect the local magnetic field environment away from the ends of the rectangular ring does not change. In other
words, the discharge plasma experiences virtually the same magnetic environment as its smaller area counterpart.
This insensitivity of the magnetic environment to lateral growth (particularly in the case of the microwave plasma
production approach) is in sharp contrast with cylindrical devices where the curvature of each magnet ring and thus
the local magnetic field changes appreciably with changes in diameter. In this regard, magnetic circuit re-design is
necessary to assure comparable performance in cylindrical devices as it is scaled up in diameter.

Ion Extraction Electrodes

Like the discharge chamber, the ion optics electrodes are also of rectangular geometry. The 2-grid ion extraction

system is manufactured from flat, pyrolytic graphite sheet. The pyrolytic graphite sheet has a significantly lower
sputter yield at relevant energies than molybdenum, which is used for the NSTAR and NEXT thrusters. For
example, the sputter yield of xenon ions on carbon is 1/5 that of xenon on molybdenum at 500 eV. Additionally, the
grids are large in cross-sectional area—over 5 times that of the NSTAR thruster. The large grid extraction surface
area allows reduction in beam current density, which in turn contributes to reduced charge exchange erosion rates.
To increase the beam extraction area of a conventional circular thruster, the ion optics diameter must increase. The
grid span to gap ratio increases proportionally, resulting in a more challenging mechanical design. With grid gap
controlled by electrical standoffs located only around the perimeter, the mid-span region may be subject to
significant gap variability. Deformation due to electrostatic attraction, thermal strain, and launch vibration become
worse as the unsupported span increases in length. This is unfortunate since typically for hollow cathode driven ion
thruster discharges, the center region often has the highest beam current density and thus highest thermal load. With
a rectangular ion optics geometry, the maximum length of unsupported span is bounded by the rectangle width. Ion
beam extraction area can be increased significantly by increasing thruster length, but the grid gap remains closely
controlled across its width.

The aforementioned ion optics attributes give the HiPEP thruster significant life margin. Indeed, the baseline

grid geometry accommodates a 100 kg/kW throughput with margin at both the 8000 s and 6000 s specific impulse
design points. Further details regarding ion optics performance may be found in reference 25.

The Neutralizer

The HiPEP neutralizer must satisfy a number of stringent requirements:

Provide up to 6-9 A of electron emission current necessary for beam neutralization.

Demonstrate growth potential electron emission currents in excess of 9 A.

The neutralizer must be capable of supplying 6-9 A continuously for periods 7-14 years

Erosion processes must be well understood and appropriately addressed by both model and
extended wear tests.

Electron extraction voltages must be sufficiently low (ideally less than threshold sputter energy).

Multiply-charged ion fractions must be minimized to reduce the erosion of neutralizer surfaces.

Electron emitting temperature and chemistry must be well understood to ensure there will be no
migration of materials to the vicinity of the cathode orifice.

The neutralizer design must attempt to minimize mass, volume and gas flow requirements with the NEXT
neutralizer as a baseline reference.

HiPEP thruster design points are optimized such that flow allotment for the neutralizer at 3.5 A
beam is approximately 5 sccm and 7 sccm at 6 A beam current. Neutralizer flow rate optimization
impacts total propellant efficiency, specific impulse, and total thruster efficiency. (A beam current
of about 3.5 A is required for operation at 25 kW and 8000 s specific impulse.)

Two different neutralizer approaches are being investigated under the HiPEP project. The baseline approach

utilizes a conventional hollow cathode. This baseline approach was selected because conventional hollow cathode-
based neutralizers are 1) capable of supplying the required electron flux at acceptable expenditure of power and
xenon flow, and 2) the neutralizer assembly undergoes reduced erosion relative to the discharge cathode.

This

NASA/TM—2004-213194

latter point is bolstered by observed state of the
ELT NSTAR thruster neutralizer determined after
termination of the wear test.

After over 30,000

hours of testing, with the exception of the
underside of the keeper tube that faces the beam,
the neutralizer’s condition was fairly pristine.
Indeed, the neutralizer continued to operate
nominally over the duration of the wear test. The
reason for the apparent immunity to degradation
resides in the fact that the neutralizer is not
immersed in a dense plasma. Additionally, the
potential of the neutralizer and the local space
potential is typically less than the sputter threshold
for neutralizer materials.

The presence of the keeper tube underside

erosion determined post ELT suggests direct
impingement or enhanced charge exchange
erosion. Fabrication of the keeper from graphite
would address enhanced charge exchange erosion
occurring between the beam plasma and the
neutralizer plasma. It, however, does not address
potential erosion driven by extreme off-axis beam
ions. This issue can be practically addressed by
optimizing the neutralizer’s position. Such an
optimization study is necessary in order to avoid off-axis direct impingement, which over test durations much longer
than the NSTAR ELT could be a potential life limiter.

The above-mentioned issues address erosion issues associated with physical sputtering. In addition to physical

sputtering, the neutralizer lifetime is also a function of emitter condition. Barium depletion for all practical purposes
represents emitter end-of-life. Provided the physical sputtering issue is solved via position optimization and the
integration of a graphite keeper, two approaches are considered to extend neutralizer system life. First, design
criteria, life models, and supporting neutralizer test data will be obtained to validate the lifetime of conventional
hollow cathodes meeting or exceeding JIMO requirements. Secondly, multiple neutralizers may be employed. The
number of neutralizers required is determined by the defined life per neutralizer. The product of the life per
neutralizer and the total number of neutralizers can be optimized such that it exceeds the mission lifetime
requirement by approximately 1.5 (for margin.)

A microwave neutralizer is also being developed under the HiPEP activity. The basis of this activity is primarily

risk mitigation. Microwave neutralizers have been used successfully for at least one deep space mission—MUSES
C.

27-28

In the strictest sense, the microwave neutralizer is essentially a plasma cathode. This concept is illustrated in

Fig. 3. Electrons are extracted from the boundary of a very dense discharge plasma.

Extracted electrons can also

generate additional electrons via collisions with gas exiting the neutralizer. This plasma bridge reduces impedance
and thereby reduces the neutralizer extraction voltage. Integration of a microwave neutralizer with a microwave
main discharge plasma generator in addition to providing extended life, also simplifies power supply and feed
system requirements and eliminates gas cleanliness protocols. Both internal antenna and slotted antenna approaches
are viable options for neutralizer plasma generation that have been actively pursued under the HiPEP project.

III.

Thruster Evolution

From a manufacturing standpoint, the rectangular prism shape of the discharge chamber and the planar

rectangular grid electrodes make fabrication fairly straightforward. Indeed, involved manufacturing processes such
as spin-forming, typical of cylindrical geometries and ion optics dishing, are not required. The discharge chamber
itself is consists of flat stainless steel panels to which flake containment mesh is attached. Assembly of the panels
using angle bracket or the like is relatively straightforward. Magnets are mounted on the outside of the rectangular
discharge chamber. The magnets are attached to the outer shell via stainless steel channel retainers. Mounting the
magnets on the outside reduces the heat load to an individual magnet ring. Integration of the plasma generation
approach, be it hollow cathode or microwave, is also straightforward.

Plasma

Electrons

Power Supply

Discharge
Chamber

Collector electrode

(physical or virtual, eg. Beam)

Plasma

Electrons

Power Supply

Discharge
Chamber

Collector electrode

(physical or virtual, eg. Beam)

Figure 3. Generalization of a plasma cathode electron
source. Electron current is extracted from a dense
plasma formed within the discharge cavity. Cavity
plasma is generated via ECR .

NASA/TM—2004-213194

The test objectives of the laboratory version of the HiPEP thruster are to map out performance and provide

insight into optimization. These goals include: 1) optimize discharge performance (discharge losses, propellant
utilization, and plasma uniformity (beam flatness), 2) optimize grid performance (perveance margin and
backstreaming limit), and 3) evaluate neutralizer performance.

Based on results of this optimization process the overall engine design will evolve by incorporating

performance improving changes to the geometry and magnetic circuit. Early on, a first generation laboratory model
was fabricated and tested. Lessons learned from this model were used to fabricate a second generation laboratory
model. The development model that will be wear-tested incorporates performance improving changes ascertained in
the second generation laboratory model. This process of building on successive generations of thrusters improves
overall design and reliability by incorporating lessons learned and establishing heritage. Thermal and structural
optimization is also incorporated during the design evolution process. These modifications are guided by actual test
data and modeling such that the end product of is a high fidelity, development model thruster.

IV.

Thruster Wear Test

A preliminary assessment of the thruster’s performance over time is planned for the HiPEP thruster. The

duration of the test will be 2000 hrs. In addition to assessing the thruster’s performance over time, another function
of the test will be to determine any previously unknown failure modes. Though the wear test duration is
considerably short relative to the required lifetime of the thruster system (~10 years), some insight regarding lifetime
is expected to be gleaned from the test. For example the issue of flake containment, neutralizer keeper and discharge
cathode keeper erosion, and unexpected grid erosion can be assessed from such a test. Findings from the wear test
will be utilized in the JPL–led JIMO thruster life evaluation task.

For wear test results to be meaningful, the wear test facility has to have sufficient pumping speed as well as

sufficiently low backsputter rate. Poor background pressure enhances charge exchange erosion and affects thruster
performance and wear. High backsputter rates give rise to deposition that could mask any erosion accumulated over
the 2000 hrs. Vacuum facility 6 at NASA GRC will be the HiPEP wear test cell. The tank measures 7.6 m by 21m
with a pumping speed of approximately 300 kl/s on xenon.

Expected backsputter rates at the 25 kW, 8000 s

specific impulse operating point should be of order 1-2 micron/kh. Baseline diagnostics to be utilized in the wear
test include cameras, beam probes and deposition sensors (quartz crystal microbalance, witness plates, pinhole
cameras).

Thruster Performance and Development Status

To date the HiPEP thruster has been operated using two plasma generation approaches. The DC plasma

generation approach was selected as the primary plasma generation approach because it offered the lowest risk to
schedule. The primary objective of the backup microwave effort was to have a fully developed, high performance
microwave plasma generator available if in the event, life limitations or implementation issues associated with
hollow cathode technology enhances risk to the project. The first HiPEP engine beam extraction test was conducted
at beam powers up to 16 kW using 2.45 GHz microwaves. The design microwave frequency for the HiPEP engine is
actually 5.85 GHz. In the early test, 2.45 GHz was used because it was available and provided a low risk
demonstration of the concept. The higher frequency operating design point significantly increases the plasma density
and thus propellant utilization as well as maximum extractable beam current. Specific impulse for 2.45 GHz test
ranged between 4500s to 5500s. The test illustrated that large volume, uniform plasma could be generated using
microwave ECR. Figure 4 illustrates microwave thruster in operation along with an ion beam profile. Beam flatness
(the ratio of average to peak current density) of over 0.82 was measured with microwave ECR, demonstrating the
ability to produce uniform plasma profiles at the ion extraction plane. The test represented the largest (size and
power) ECR ion source ever operated. Subsequent microwave engine testing was done at 5.85 GHz, the thruster
design frequency. These higher frequency tests were aimed at discharge performance optimization. Microwave
thruster discharge testing at 5.85 GHz demonstrated the higher plasma production capacity as compared to
2.45 GHz. At this higher frequency, simulated ion grid currents over 4 A were measured, well above that which is
needed for the 8000 s design point.

The DC HiPEP engine has been performance characterized at power levels up to 40 kW. A photograph of the

engine operating at 34 kW is shown in Fig. 5. Table I presents typical thruster performance over a power range
between 10 and 40 kW. The data shown in the table has been corrected for ingested flow. The nominal specific
impulse is approximately 8000 s, the primary design point. As can be seen here, there thruster performs well over a
specific impulse range between 6000 and 10000 s. Total thruster efficiencies in excess of 75% were achievable at
the higher power levels as indicated in the table.

NASA/TM—2004-213194

Table I. DC HiPEP Thruster Performance

Power, kW

Flow Rate,
mg/s

Efficiency

Thrust,
mN

Specific
Impulse

9.7

4.0

0.72

240

5970

15.9

4.9

0.74

340

7020

20.2

5.6

0.75

410

7500

24.4

5.6

0.76

460

8270

29.6

6.2

0.80

540

8900

34.6

6.6

0.77

600

9150

39.3

7.0

0.80

670

9620

(a)

(b)

Figure 4. Photograph of HiPEP engine operating with microwave ECR as the plasma production
approach. (b) Beam profile at a beam current of 1.64 A.

NASA/TM—2004-213194

Discharge chamber performance

was assessed for the 8000 s specific
impulse design point by plotting the
discharge losses versus the discharge
propellant utilization efficiency as
shown in Fig. 6. Here the beam current
and discharge voltage were fixed as
flow rates and discharge current are
adjusted to vary the discharge
propellant utilization efficiency.
Discharge losses were plotted as a
function of utilization efficiency for
two different discharge voltages. The
discharge voltage at a given internal
discharge chamber pressure determines
the nature of the electron energy
distribution function. In this respect,
the discharge voltage will affect
ionization efficiency for a given input
total flow. The doubly charged xenon
fraction typically increases at high
propellant utilizations and associated
high discharge losses. It is desirable to
operate in the knee of the utilization
curve because it is here where the
required discharge power expenditure
for a given beam current is minimized.
It was observed that beyond the knee,
at the higher propellant utilization
efficiencies (>0.85), discharge losses
were lower at the higher discharge
voltage. This is likely related to
improved ionization efficiency at the
higher discharge voltage. For example,
at 28 V, the discharge losses for
operation at propellant efficiencies of
0.90 and 0.92 are 188 and 196 W/A,
respectively. This is to be contrasted
with the 25 V data in which discharge
losses were greater 200 W/A for
propellant efficiencies greater than
0.90. Because discharge power is a
small fraction of total thruster power, operating at the 28 V discharge does not significantly impact performance. An
assessment of the fraction of doubly charge ions will be necessary to determine if erosion is an issue for operation at
28 V.

Figure 7 illustrates the range and growth capacity for the ion thruster. Here thruster efficiency and specific

impulse are plotted as a function of thruster power. For each curve, the propellant utilization efficiency and beam
current were held fixed. The beam voltage was then varied to throttle in power. In general, under the conditions of
fixed beam current and utilization, the specific impulse should increase as the square-root of the thruster power,
provided the discharge power is a small fraction of the total input power (a situation which prevails here.) As can be
seen here, the specific impulse increases as expected monotonically with increasing power (beam voltage). To verify
the square-root dependence, a function of the form

⋅

, where a and b are fitting parameters, was fit to each

power throttling curve. The constant b should be approximately 0.5, indicating the expected square root relationship
with power. For the data shown here, b was found to be approximately 0.53, which is in good agreement with the
expected scaling.

0.70

0.75

0.80

0.85

0.90

0.95

1.00

160

170

180

190

200

210

220

230

240

Discharge Voltage = 25 V
Discharge Voltage = 28 V

Dis

charg

ssess,

Discharge Propellant Utilization Efficiency

Figure 6. HiPEP engine discharge losses at the 8000 s design point.

Figure 5. Photograph of the DC HiPEP engine operating at 34 kW.

NASA/TM—2004-213194

Thruster efficiencies greater than

65 % (NRA requirement) were
demonstrated even for power levels
as low as 6 kW as indicated in Fig. 7.
Here, the solid line indicates the
NRA efficiency requirement. The
thruster efficiency indicates a small
positive slope as beam voltage
(thruster power) is increased. Strictly
speaking, if ion extraction efficiency
does not vary, the thruster efficiency
should be essentially flat or constant
over the power range. The small
deviation is most likely attributed to
an increase in screen grid
transparency which typically
increases with increasing beam
voltage.

This increase results in a

slightly reduced discharge power
requirement for a given beam current,
thereby resulting in an improvement
in thruster efficiency with increasing
thruster power. Power levels up to
nearly 40 kW are also illustrated in
Fig. 7. Power levels above 40 kW
were limited only by power supply
output capacity. In all cases,
discharge propellant utilization
efficiency was ~0.9 or better. Over
this power range, specific impulse
values of 7000 s to 9000 s were
demonstrated. The family of curves
in Fig. 7 illustrates the growth
potential and overall range of the
thruster.

Thrust as a function of input

power at a nominal specific impulse
centered at 8000 s is illustrated in
Fig. 8. The beam voltage is
essentially fixed for the data
presented in the figure. Under these
conditions, thrust should be a linear
function of thruster input power. This
functional relationship is illustrated
in the figure. For all data points
illustrated in the figure, the average
thruster efficiency was greater

than 75%.

Presently, a development model thruster is being prepared for wear testing. It will be nearly identical to the

previous generation thruster used in performance tests with a few notable exceptions: 1) magnet rings will be
attached to the outside of the discharge chamber 2) flake containment mesh will be installed on discharge chamber
inner surfaces 3) ion optics mount system will be of higher fidelity from a structural standpoint, and 4) the discharge
cathode will feature a graphite keeper electrode. The objective of the wear test will be to measure the thruster’s
capacity to operate reliably over extended duration, reveal yet unidentified failure mechanisms, and make a limited
assessment on thruster life.

300

400

500

600

700

800

Specific Impulse ~ 8000 s

Thru

st, mN

Thruster Input Power, kW

Figure 8. HiPEP thruster thrust variations as a function of thruster
input power.

0 2 4 6 8 10 12 14 16 18 20 22 24 26 28 30 32 34 36 38 40

5000

6000

7000

8000

9000

10000

Thruster Power, kW

Specif

Impulse,

Specific Impulse

0.0

0.2

0.4

0.6

0.8

1.0

NRA efficiency requirement

Th
ru

er Ef

fici

en
cy

Thruster Efficiency

Figure 7. HiPEP engine power throttling range demonstrates high
efficiency over a wide power range.

NASA/TM—2004-213194

VI.

Concluding Remarks

The HiPEP project was one of three research efforts selected to develop a high power electric propulsion system

to satisfy thruster requirements for NEP missions, in particular the proposed JIMO mission. The HiPEP project
selected a large area, rectangular thruster geometry to address the requirements of high power and long life. The
HiPEP project thruster has been fabricated and tested. The engine’s design includes the integration of technologies
aimed at increasing life well above to the state of the art by a factor of 3. These technologies include rectangular,
pyrolytic graphite grids, large area discharge chamber and a discharge cathode assembly with a graphite keeper
electrode. Microwave ECR plasma generation technologies are also being developed to mitigate risk. Performance
testing confirms thruster growth potential to power levels well beyond the targeted 25 kW operating point.
Efficiencies in excess of 0.75 have been measured over a range of thruster powers (20-40 kW). Presently a HiPEP
development model thruster is being prepared for a 2000 hour wear test. The objective of the wear test is to
demonstrate reliable extended duration operation as well as provide insight into wear mechanisms.

References

El-Genk, M.S., ed., A Critical Review of Space Nuclear Propulsion: 1984-1993, AIP Press, NY, 1994, pp. 22-23.

Sengupta, A., et al, “Status of the Extended Life Test of the Deep Space 1 Flight Spare Ion Engine after 30,000 hours of

Operation,” AIAA Paper 2003-4558, July 2003.

Patterson, M.J., Roman, R.F., and Foster, J.E., "Ion Engine Development for Interstellar Precursor Missions," AIAA Paper

2000-3811, July 2000.

Oleson, S., “Electric Propulsion Technology Development for the Jupiter Icy Moon Orbiter Project,” AIAA Paper 2004-

3449, July 2004.

Rawlin, V.K, Williams, G.W., Pinero, L.R., and Roman, R. F., “Status of Ion Engine Development for High Power, High

Specific Impulse Missions,” IEPC Paper 2001-096, October 2001.

Polk, J.E, et al, “An Overview of the NEXIS Program,” AIAA Paper 2003-4713, July 2003.

Patterson, M.J., Foster, J.E., Haag, T., Soulas, G.C., Pastel, M.R., and Roman, R.F.,“NEXT: NASA’s Evolutionary Xenon

Thruster Development Status,” AIAA Paper 2003-4862, July 2003.

Foster, J.E. et al, “An Overview of the HiPEP Project,” AIAA Paper 2004-3453, July 2004.

Kovaleski, S.D., Patterson, M.J., Soulas, G.C., and Verhey, T.R., “A Review of Hollow Cathode Testing for the

International Space Station Plasma Contactor,” IEPC Paper 01-271, October 2000.

Foster, J.E., Roman, R., Soulas, G.S., and Patterson, M.J., “Magnetic Grid for Electron Backstreaming Mitigation,” IEPC

Paper 01-221, October 2001.

Soulas, G.C., Foster, J.E., and Patterson, M.J., “Performance of Titanium Optics on a NASA 30 cm Ion Thruster,” AIAA

Paper 2000-3814, July 2000.

Shotwell, R., “Carbon-Carbon Grid Development for Ion Propulsion Systems,” IEPC Paper 2001-093, October 2001.

Haag, T., Patterson, M.J., and Soulas, G.C., “Carbon-Based Ion Optics Development at NASA GRC,” IEPC Paper 2001-

094, October 2001.

Soulas, G.C., “Improving Total Impulse Capability of the NSTAR Ion Thruster with Thick-Accelerator –Grid Ion Optics,

IEPC Paper 01-081, October 2001.

Yamamura, Y. and Tawara, H., “Energy Dependence of Ion-induced Sputtering Yields from Monatomic Solids at Normal

Incidence,” Atomic Data and Nuclear Data Tables , Vol. 62, 149-253, 1996.

Foster, J.E. and Patterson, M.J., “Discharge Characterization of 40 cm-Microwave ECR Ion Source and Neutralizer,” AIAA

Paper 2002– 5012, July 2002.

Divergilio, W.F., Goede, H., and Fosnight, V.V., “High Frequency Plasma Generators,” NASA CR-167957, November

1981.

Toki, H., Fujita, H., Nishiyama, K., Hitoshi, K., and Funaki, I., “Performance Test of Various Discharge Configurations for

ECR Discharge Ion Thruster,” IEPC Paper 01-107, October 2001.

Kuninaka, H., Satori, S., Funaki, I., Shimizu, Y., and Toki, K., “Endurance Test of Microwave Discharge Ion Thruster

System for Asteroid Sample Return Mission MUSES-C,” IEPC Paper 92-137, December 1997.

Toki, K., et al., “Technological Readiness of Microwave Ion Engine System for MUSES-C Mission,” IEPC 01-174,

October 2001.

Park, S.S. et al., “Reliability Analysis of Klystron-Modulator System,”

Proceedings of the Second Asian Particle

Accelerator Conference

, Beijing, China, 2001.

Hansen, J.W., “US TWTs from 1 to 100 GHz,”

Microwave Journal

, 1989, Vol. 32, pp. 179-183.

Dieumegard, D. et al, “Life Test Performance of Thermionic Cathodes,”

Applied Surface Science

, Vol. 111, February,

1997, pp. 84-89.

Sovey, J.S., “Improved Ion Containment using a Ring-Cusp Ion Thruster,”

Journal of Spacecraft and Rockets

, Vol. 21,

Sept.-Oct. 1984, pp. 488-495.

Williams, G.W. , Sovey, J., and Haag, T., “Performance Characterization of HiPEP Ion Optics,” AIAA Paper 2004-3627,

July 2004.

NASA/TM—2004-213194

Sengupta, A., Brophy, J.R., Garner, C.E., and Anderson, J.A., “An Overview of the Results from the 30,000 Hour Life Test

of the Deep Space One Flight Spare Ion Engine,” AIAA Paper 2004-3608, July 2004.

Toki, K, Kuninaka, H., Nishiyama, K., Shimizu, Y., and Funaki, I., “Technological Readiness of Microwave Ion Engine

system for Muses-C mission,” IEPC 01-120, October 2001.

Satori, S, Kuninaka, H., Ohtaki, M, and Ishikawa, Y., “Operational Characteristics of Microwave Discharge Neutralizer,”

IEPC Paper 97-05, December 1997.

Oaks, E. M., “ Physics and Techniques of Plasma Electron Sources,”

Plasma Sources Sci. Technol

., Vol. 1, 1992, pp. 249-

255.

Kamhawi, H., Foster, J., and Patterson, M.J., “Parametric investigation of an ECR cathode,” AIAA Paper 2004-3819, July

2004.

31”

ElectricPropulsionLaboratory:Facility Capabilities,” http://facilities.grc.nasa.gov/epl/epl_caps.html.

Soulas, G.C., Domonkos, M.T., and Patterson, M.J., “Performance Evaluation of the NEXT Ion Engine,” AIAA Paper

2003-5278, July 2003.

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

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NASA TM—2004-213194
AIAA–2004–3812

E–14693

WBS–22–982–10–02

The High Power Electric Propulsion (HiPEP) Ion Thruster

John E. Foster, Tom Haag, Michael Patterson, George J. Williams, Jr.,
James S. Sovey, Christian Carpenter, Hani Kamhawi, Shane Malone,
and Fred Elliot

Nuclear electric propulsion; Ion thruster; Electric propulsion; Specific impulse; Microwave
plasma; Hollow cathode

Unclassified - Unlimited
Subject Category: 20

Distribution: Nonstandard

Prepared for the 40th Joint Propulsion Conference and Exhibit cosponsored by the AIAA, ASME, SAE, and ASEE, Fort
Lauderdale, Florida, July 11–14, 2004. John E. Foster, Tom Haag, Michael Patterson, Hani Kamhawi, Shane Malone, and
Fred Elliot, NASA Glenn Research Center; George J. Williams, Jr., Ohio Aerospace Institute, Brook Park, Ohio 44142;
James S. Sovey, Alpha-Port, Inc., Cleveland, Ohio 44135; and Christian Carpenter, QSS Group, Inc., Cleveland, Ohio
44135. Responsible person, John E. Foster, organization code 5430, 216–433–6131.

Practical implementation of the proposed Jupiter Icy Moon Orbiter (JIMO) mission, which would require a total delta V
of approximately 38 km/s, will require the development of a high power, high specific impulse propulsion system. Initial
analyses show that high power gridded ion thrusters could satisfy JIMO mission requirements. A NASA GRC-led team
is developing a large area, high specific impulse, nominally 25 kW ion thruster to satisfy both the performance and the
lifetime requirements for this proposed mission. The design philosophy and development status as well as a thruster
performance assessment are presented.