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Field Emission Electric Propulsion:
Enabling future Science and Earth Observation Missions
Bernhard Seifert, Laura Bettiol,
Nembo Buldrini, Martin Eizinger
FOTEC Forschungs- und
Viktor Kaplan-Strasse 2
Wiener Neustadt, 2700 Austria
Viktor Kaplan-Strasse 2
Wiener Neustadt, 2700 Austria
Jose Gonzalez Del Amo, Luca Massotti
European Space Agency
Noordwijk, 2200 AG, The Netherlands
Abstract— The Field Emission Electric Propulsion (FEEP) tech-
nology based on Liquid Metal Ion Sources (LMIS) has been de-
veloped for over two decades by FOTEC Forschungs- und Tech-
nologietransfer GmbH, the research subsidiary of the Univer-
sity of Applied Sciences in Wiener Neustadt, Austria. Between
2015 and 2017 FOTEC developed the IFM NANO Thruster, an
integrated Electric Propulsion (EP) system for small satellites
and CubeSats, that was successfully in-orbit tested in coopera-
tion with FOTEC’s spin-out ENPULSION GmbH in 2018. Its
main application fields are orbit injection, drag compensation,
orbit raising, formation flight and controlled de-orbiting of
small satellites as well as very precise attitude control for larger
spacecraft. Later the system was fully qualified and commer-
cialized by ENPULSION.
Upcoming science and Earth observation missions can strongly
benefit from such a mature technology: one application example
is the ESA/NASA joint Next Generation Gravity Mission
(NGGM) where highest thrust stability and reliability are of ut-
most importance to allow precise formation flight of two space-
craft measuring the Earth gravity field at an unprecedented ac-
Compared to other electric propulsion technologies, the FEEP
thrusters offer a very dynamic thrust range from lower than 1
µN up to 1 mN. The Power-to-Thrust Ratio (PTR) can be ad-
justed during operation to allow a trade-off between available
electric power and propellant consumption efficiency. In addi-
tion, the used metallic propellant brings the great advantage of
not having pressurized vessels, valves, flow controllers and feed
lines, and indium is cheap compared to xenon, which is com-
monly used on Gridded Ion Thrusters (GIT) and Radiofre-
quency Ion Thrusters (RIT) that operate in a similar thrust
FOTEC has conducted an endurance test of a single FEEP emit-
ter for more than 50,000 hours of accumulated firing time. This
test has proven the stability and suitability of this technology for
highly demanding missions, such as NGGM, where more than
six years of continuous in-orbit operation are required. The per-
formance of FOTEC’s FEEP thruster was characterized with
the use of a thrust balance and an advanced plasma diagnostic
system consisting of 23 Digital Faraday Cups (DFC) and a Re-
tarding Potential Analyzer (RPA). This allowed comparison be-
tween direct and indirect thrust measurement and the verifica-
tion of the performance model.
Recently an ion-lensing system has been developed to reduce the
plume cone of the thrusters and to increase thrust level for the
same beam power. FOTEC is currently developing a new gen-
eration of FEEP thrusters that will provide an increased total
impulse, will require lower heating power and will offer in-
creased mass efficiency and longer lifetime.
This paper will not only give an overview and advantages of the
FEEP technology in general, but also reveal acquired test results
gathered in recent years. The current development and qualifi-
cation efforts that will allow use of FOTEC’s proprietary FEEP
technology for highly demanding future science and Earth ob-
servation missions are highlighted.
TABLE OF CONTENTS
1. INTRODUCTION ....................................................... 1
2. FEEP THRUSTERS OVERVIEW .............................. 2
3. PERFORMANCE AND ENDURANCE TESTING .......... 3
4. RECENT DEVELOPMENTS ....................................... 5
5. MANEUVERS ............................................................ 6
6. IN-ORBIT EXPERIENCE ........................................... 8
7. DEVELOPMENT ROADMAP .................................... 8
8. CONCLUSIONS ........................................................ 9
ACKNOWLEDGEMENTS .............................................. 9
REFERENCES ............................................................... 9
BIOGRAPHY ............................................................... 11
Liquid Metal Ion Sources (LMIS) and Field Emission Elec-
tric Propulsion (FEEP) are two promising technologies in the
field of advanced propulsion systems. These technologies
represent cutting-edge advancements in the quest for more
efficient and precise propulsion methods for spacecraft.
Among the key players in the development of FEEP thrusters,
FOTEC Forschungs- und Technologietransfer GmbH
(FOTEC), the research subsidiary of the University of Ap-
plied Sciences in Wiener Neustadt (Austria) with its spin-out
ENPULSION GmbH have emerged as notable contributors.
This introduction will delve into the fundamentals of LMIS
and FEEP while highlighting FOTEC and ENPULSION’s
significant efforts in advancing FEEP thruster technology.
LMIS are devices used to generate and emit ions for various
applications, including ion propulsion systems for spacecraft.
The basic principle behind LMIS involves the use of a liquid
metal which is ionized and emitted as a focused beam. This
ion beam can be accelerated and expelled to produce thrust,
enabling precise spacecraft maneuvering and long-duration
missions. LMIS technology has evolved over the years, of-
fering higher efficiency and performance compared to other
electric propulsion systems or traditional chemical propul-
sion systems. FOTEC has a long heritage in flying LMIS
technology: the first demonstration of the technology in orbit
was on the space station Mir at the beginning of the 1990s
, up to the Magnetospheric Multiscale Mission (MMS) of
the National Aeronautics and Space Administration (NASA)
, launched in 2015.
FEEP represents a subset of ion propulsion, where ions are
emitted through field emission from a sharp metallic emitter,
like the one shown in Figure 1.
Figure 1. FEEP emitter with propellant reservoir.
FEEP thrusters are known for their exceptional precision and
high specific impulse, making them ideal for applications re-
quiring fine control and extended missions. More details on
the functioning principle, operational envelope and ad-
vantages compared to other electric propulsion technologies
are given in section 2.
Over the last years, several projects have been carried out by
FOTEC with the support of the European Space Agency
(ESA), in order to optimize the FEEP technology to be used,
e.g., on the Next Generation Gravity Mission (NGGM) ,
, a joint mission with NASA, where extremely precise
thrust control is required to fulfill the mission objectives. In
addition, the mission profile of this Earth observation mission
requires long duration reliability (lifetime requirement: 5.5
years, goal: 7.5 years). During these research projects, per-
formance testing and endurance testing of the FEEP technol-
ogy have been carried out, and an overview of the results is
given in section 3. Moreover, section 4 shows the ongoing
developments that are being done to mitigate possible failure
modes and to improve the performance, respectively by using
a finned extractor and a focus electrode. The latter helps to
minimize the cosine losses by reducing the plume divergence.
The first in-orbit demonstration of a FEEP-based integrated
propulsion system was successfully done in 2018 on a IFM
NANO Thruster . Since then, the FOTEC’s spin-out com-
pany ENPULSION has launched 170 thrusters – for
different purposes: from attitude control to orbit raising up to
Geostationary Earth Orbit (GEO). More details on applica-
tions and maneuvers that are possible using FEEP thrusters
are given in section 5, while section 6 gives details on the
current status of the FEEP technology in orbit. Section 7 con-
cludes the paper with an overview of the future developments
and a summary is given in section 8.
2. FEEP THRUSTERS OVERVIEW
FEEP thrusters are known for their high specific impulse,
which can be as high as 10,000 s. This makes them suitable
for applications where fuel efficiency is critical, such as deep
space or long-duration missions. At nominal operation, a spe-
cific impulse of 2,000 to 9,000 s can be achieved . Typi-
cally, FEEP systems deliver thrust in the µN to mN range,
with a thrust resolution below 1 µN. FOTEC and EN-
PULSION use indium as propellant, which has a low vapor
pressure at typical operating temperatures, a low work func-
tion making it well suited for field emission, as well as high
density, reducing the footprint for a given total impulse. In
addition, it is non-toxic, and solid in ambient conditions,
making it easy to handle on ground and safe during launch.
Nonetheless, the melting temperature is relatively low (156.6
°C). The FEEP technology was particularly optimized for the
Laser Interferometer Space Antenna (LISA) Pathfinder mis-
sion as backup technology –. Later, the focus was
laid on porous tungsten crowns instead of single needle emit-
The functioning principle is shown in Figure 2. By applying
a high positive electric potential to the emitter and a high neg-
ative electric potential to the extractor electrode, propellant
atoms are ionized, extracted, and accelerated, resulting in a
gentle yet efficient thrust.
Figure 2. FEEP functioning principle .
The thrust being generated with FEEP thrusters can be calcu-
lated as follows:
with the emitted current (difference between emitter cur-
rent and extractor current), the emitter voltage , the mass
of an indium atom , the elementary charge and the thrust
coefficient . This value, also called divergence efficiency
coefficient, mainly describes the cosine losses due to the di-
vergence of the beam.
The specific impulse can be expressed as
where is the gravitational acceleration and is the mass
efficiency, a coefficient that describes the ratio of ionized
propellant and the total mass loss over a certain period of
time. The experience with FEEP technology gathered in the
last decades led to the development of the first integrated pro-
pulsion system based on FOTEC’s proprietary technology.
From 2015 to 2017 the IFM NANO Thruster (see Figure 3)
was developed and tested. Details on the performance can be
found in –.
The integrated propulsion system consists of the porous tung-
sten crown emitter, thermionic neutralizers for space charge
compensation and integrated power electronics to control the
thruster subsystems and to acquire telemetry. A variety of test
campaigns allowed to characterize the performance of the
thruster at different operation points.
Figure 3. The first integrated propulsion system built on FEEP
technology, the IFM NANO Thruster.
3. PERFORMANCE AND ENDURANCE TESTING
The primary performance metrics of a FEEP engine are
thrust, specific impulse, thrust vector, and divergence angle.
These data require different methods of testing, specifically,
for example, direct thrust measurement (DTM), a retarding
potential analyzer (RPA) to measure the energy distribution
of the ion beam and thus the exhaust velocity, and Faraday
cups (FCs) to measure the ion current distribution of the beam
and thus the beam divergence and the direction of thrust.
The aforementioned testing equipment was developed and
built in-house to maximize accuracy, reduce unknown error
sources, match data acquisition and post-processing, and to
be able to take into account technology-specific considera-
tions such as the propellant species and operating ranges al-
ready in the design phase.
DTM was performed for different thrusters that use FOTEC’s
proprietary FEEP technology, at several operating points.
Denoising is performed by the use of an exponential moving
average (EMA) filter. Thrust balance drift/offset are ac-
counted for in the presented data by taking time-averages of
short time windows after thrust is turned off and subtracting
that from the smoothed raw data of the preceding measure-
ment window. Figure 4 shows the offset-corrected thrust bal-
ance data, and the FEEP system thrust setpoint and reading.
Figure 4. Direct thrust measurements of a FEEP thruster.
Previous measurements  of the ion energy distribution of
a FEEP system using a 4-grid RPA located at a distance of
0.95 m from the point of emission indicate that most ions
have an energy equal to one elementary charge times the
emitter voltage (see Figure 5).
Figure 5. Ion current and energy density in the beam center.
It was furthermore determined that the ion energy does not
depend on the direction in which the ion is accelerated, as is
shown in Figure 6. This strongly supports the approach of
computing the thrust vector from many small thrust vector
contributions computed from the ion current density at many
individual points. Otherwise, the ion energy at each point
would also need to be measured.
Thrust vector measurements were performed using spatially
distributed current density measurements at 23 vertical and
160 horizontal angles, where the horizontal positions are
sampled successively. The data acquisition was performed
with Digital Faraday Cups (DFCs) with internal signal con-
ditioning, high Signal-to-Noise Ratio (SNR) and 24-bit A/D
conversion resolution with less than 2% total error –.
By assuming linear trajectories and isotropic ion energy dis-
tribution, the local ion current is linearly related to thrust, al-
lowing for calculating the thrust contribution of each meas-
urement point. Using this method, which proved outstanding
correspondence to direct thrust measurement , , the
beam profile of two juxtaposed ions sources operating simul-
taneously with their centers of emission approximately 10 cm
apart could be obtained (see Figure 7).
Figure 6. Angular mapping of the ion energy.
Figure 7. Current density, thrust vector, and divergence, for
the superposition of two FEEP ion beams.
Figure 8 shows the measured ion current density of a similar
setup but with an electrical lensing system attached to both
thrusters. This lensing system (also known as focus) is pre-
sented in section 4.
Figure 8. Current density, thrust vector, and divergence, for
the superposition of two focused FEEP ion beams.
Similar beam diagnostics measurements were performed at
an hourly rate over a duration of 250 hours. The computed
vertical and horizontal thrust vector angles are shown as
and in Figure 9.
Figure 9. Long-term measurements of thrust vector misalign-
ment from plasma diagnostics.
In addition to performance tests at beginning-of-life (BOL),
multiple endurance tests were carried out on a porous crown
emitter. The initial test was conducted for 10,000 hours ,
approximately corresponding to the depletion of the 250 g
tanks with a mass efficiency of 50%. Following the good per-
formances of the emitter, the test was extended in multiple
steps, until reaching the final milestone of 48,000 hours. At
the end of the 48,000 hours, after examination of the emitter,
it has been decided to further extend the operation of the
thruster until reaching a total operation of exactly 6 years
(52,560 hours), which was achieved very recently . The
test was conducted on a thruster module equipped with focus-
ing electrode, which underwent several maintenance steps
during the long emitter testing period, since it was initially
designed for 10,000 hours lifetime.
The emitter, on the other hand, demonstrated extreme robust-
ness and minimal modification of the tungsten substrate, even
well beyond the initially planned firing period.
Figure 10. Firing duration at different thrust levels.
Figure 10 shows how long the thruster has been operated at
specific thrust levels, for thrust values higher than or equal to
50 µN (please note that the ordinate is shown in logarithmic
scale). The two modes of 350 µN and 60 µN stand out,
corresponding respectively to the first large testing period
spanning from February 2015 to August 2020, and to the last
testing period from December 2021 to September 2023, when
the thruster was operated at reduced thrust levels.
Counting the hours for thrust values higher than or equal to
50 µN, the resulting total firing time amounted to 52,279
hours. This is shown in Figure 11. The total impulse produced
corresponded to 47 kNs. The average specific impulse (Isp)
was approximately 6,500 s, and the average mass efficiency
of the emitter was 45%.
Figure 11. Firing duration at different thrust levels.
4. RECENT DEVELOPMENTS
The results presented in the previous section extend the ex-
isting data available for FOTEC’s proprietary FEEP technol-
ogy and thereby underline the competitiveness of this tech-
nology compared to other EP solutions. In this section vari-
ous modifications developed in recent years are presented.
This includes improvements upon the baseline in many im-
portant metrics such as PTR, lifetime, thrust stability, and
thrust vector control.
The beam divergence losses can be reduced substantially by
electrostatic redirection of ions that have a large lateral ve-
locity component, schematically shown in Figure 12.
Figure 12. Electrostatic redirection of laterally moving ions.
This effect was achieved by slightly modifying an existing,
proven design and adding a pair of electrodes . The re-
sulting assembly is shown in Figure 13 .
Figure 13. Focus module on top of a FEEP thruster
The focus assembly height is ca. 4 cm and allows a net thrust
increase of ca. 30% (see Figure 14) without an appreciable
increase in power because the total current is the same. These
values were measured experimentally by means of plasma di-
agnostics similarly to that described in section 3.
Figure 14. Thrust of a FEEP system with and without lensing
system from simulation (transparent) and experiment (full
An additional benefit of the focus module is that the high
electric potentials near the emitter are screened from the sur-
roundings, making the assembly more compatible with other
electrostatically sensitive devices in the vicinity.
Mitigation to Extractor Clogging
The operation of an indium FEEP emitter produces, in addi-
tion to the ion beam, a spray of neutral indium droplets. These
droplets deposit on all the surfaces in the field of view of the
needle tips. The extractor is the thruster element which is
mostly impacted by this deposition. In the long-term
operation of the thruster, the indium depositing on the extrac-
tor can lead to an electrical connection between the emitter
and extractor, inhibiting the ability for ion emission if not re-
moved by cleaning the extractor. Mitigation measures, such
as the cleaning procedure during which indium deposition is
melted and removed from the extractor, are therefore of pri-
mary importance in order to allow the thruster operation until
the ultimate depletion of the propellant reservoir.
One possibility is to design an extractor able to store the de-
posited indium away from the emitter. A new even more ef-
ficient extractor design was conceived as a series of parallel
plates. The idea is to take advantage of capillary forces be-
tween the plates, which will be able to direct and keep in
place the molten indium, avoiding an accumulation on the
rim of the extractor. As the indium accumulates on the edges
facing the emitter, the distance between the emitter and ex-
tractor decreases. The solution is to melt the indium on the
extractor well before reaching this critical point.
On a normal extractor, when testing the thruster in a ground-
based lab for a longer time, the gravity force causes the mol-
ten indium to coalesce into a large single droplet at the bottom
of the extractor.
Using an extractor constituted by parallel plates, the liquid
indium will be absorbed inside the plates by capillary forces,
with a consequent restoration of the original distance between
the emitter and the extractor edge. This kind of extractor was
tested extensively in a test campaign where two emitters were
fired for 8,000 hours at 200 µN (see Figure 15) and 100 µN.
Figure 15. Finned extractor showing the absorption of indium
inside the interspaces.
During regular testing operations on ground, the extractor
must be cleaned at regular intervals, as indium accumulates
on the bottom side due to gravity; this entails several thermal
and vacuum-air-vacuum cycles, which have been shown to
compromise the emitter performance. The implementation of
the new type of extractor allowed operation for extended pe-
riods of time, thus eliminating thermal cycling and exposure
A different modification of the extractor, designed to provide
purely electrical steerability of the thrust vector, has also been
studied. By using three extractor segments (see Figure 16)
supplied by dedicated electronics, instead of one continuous
ring, the electric field can be varied locally and asymmetri-
cally. This method for thrust steering has been patented in
Figure 16. Segmented extractor around the crown emitter.
A model of the ENPULSION AR³  incorporating this
technology was tested at the ESA Propulsion Laboratory,
where the inclination of the ion beam was measured by means
of Faraday cups for several thrust vector setpoints .
Throughout the course of a typical mission, propulsion sys-
tems are required for different purposes. If a satellite is
launched as piggy-back, it needs to inject itself into the right
orbit after deployment. On the one hand, this adds flexibility
with respect to the available options for launches and helps to
ultimately reduce the launch costs. On the other hand, the
complexity of the satellite is increased by integration and
later operation of a propulsion system. Depending on the or-
bital elements to be changed, the required delta-v can be con-
siderable - in particular for inclination changes. For these ma-
neuvers, a propulsion system with high specific impulse and
total impulse is advantageous.
In the first example, a satellite is orbiting the Earth on a cir-
cular orbit with an initial altitude . For the change of the
altitude , the following delta-v approximation for
is required :
with the standard gravitational parameter the
Earth mass and radius . The delta-v can be expressed
as total impulse per mass with the thrust maneuver dura-
tion and spacecraft mass . Under the assumption of a 3U
CubeSat with (configuration S1) and an initial al-
titude of , a change to requires
. This corresponds to a continuous thruster operation
of at which is the nominal thrust
level of the IFM NANO Thruster. At this operation point, the
system power consumption is 40 W .
In the following, an inclination change maneuver shall be
considered. To change the plane by , the following delta-v
is required :
Considering the same spacecraft and same initial orbit with
, a change of requires
. Although for this maneuver no continuous thrust can be
applied, this velocity change would correspond to a thruster
operation time of at .
This example shows how “expensive” an inclination change
maneuver is compared to an orbital raise. Details on these and
other maneuvers can be found in .
During operation, the orbit needs to be maintained and the
Attitude and Orbit Control System (AOCS) needs to, e.g.,
compensate for perturbations and atmospheric drag. The lat-
ter is particularly important if the satellite is equipped with
large solar arrays or if the altitude is particularly low - such
as for Very Low Earth Orbits (V-LEO) which will play an
important role in the near future of the commercial satellite
sector as it provides low latency communication between the
ground station and the spacecraft.
The atmospheric drag strongly depends on the altitude and
the solar activity. With the use of the developed CubeSim
simulation framework , different scenarios were consid-
ered to compute the required thrust to counteract the drag
force. The altitudes of , and and the
timeframe from 1989 to 2023 were used.
Figure 17 shows the simulation results, whereas the clear re-
lation between the solar activity (strongly tied to the number
of sunspots) and the atmospheric drag can be recognized. For
altitude, the drag force varies between and
depending on the solar activity. For al-
titude, the drag force is between and and
for between and .
Figure 17. Atmospheric drag for different altitudes.
A typical 3U CubeSat with body-mounted solar cells has a
max. projected area of . At an altitude of , the
drag force therefore varies between and and at
between 0.06 and . These forces can be com-
pensated using an electric thruster, such as the IFM NANO.
Considering a constellation of two or more spacecraft, the
distance between the spacecraft or the orbital coverage needs
to be maintained to fulfill the intended service or science
In addition to configuration S1, two more spacecraft are con-
sidered for the following simulations: S2 is a 12U CubeSat
with and a max. projected area of and is
equipped with two IFM NANO Thrusters providing up to
, and S3 is a micro satellite with and deployable
solar arrays. It spans an area of up to and uses two
IFM MICRO Thrusters  for propulsion proving up to
In the following example, two of the defined spacecraft shall
achieve an orbital separation of within 24 hours after
ejection from the same launch vehicle. Different thrust pro-
files are applied and the separation vs. time are shown in Fig-
Figure 18. Separation of different spacecraft after ejection.
Until recently, it had to be guaranteed that the satellite deor-
bited within 25 years after End-of-Life (EOL) either pas-
sively or by using a propulsion system. In order to reduce
space debris, the USA recently announced to reduce the pe-
riod from 25 to 5 years  and ESA has since then joined
this directive with the “zero debris” agreement .
In the following example, the natural decay of the defined
satellites is simulated with CubeSim. An initial altitude of
and different start dates are used. The results are
shown in Figure 19.
While at high solar activity, natural decay occurs within 3
years, at low solar activity it can take up to 14 years for the
same satellite to deorbit. This enforces the need of propulsion
systems for the majority of small satellites launched into or-
bits with an altitude of 500 km or above .
1990 1995 2000 2005 2010 2015 2020
Drag Force [N/m²]
0 5 10 15 20 25
Spacecraft Separation [km]
Figure 19. Natural decay for different spacecraft.
6. IN-ORBIT EXPERIENCE
In 2017, The IFM NANO Thruster prototype was integrated
into a 3U spacecraft with a total launch mass of 4.8 kg. Re-
action wheels ensure proper alignment of the spacecraft. The
satellite was launched January 12, 2018 on PSLV-C40 and
was released into a 491 x 510 km sun-synchronous orbit. Af-
ter commissioning, two orbit raise maneuvers were per-
formed. For the first maneuver the thruster was fired at
for 15 minutes, which resulted in an anticipated altitude
change of and was confirmed with GPS telemetry
( ). For the second maneuver, the thruster was used
for 30 minutes at the same thrust level. Due to saturation ef-
fects of the used reaction wheels, the expected altitude
change was only , which was confirmed as well
( ). Details can be found in  and .
To date, over 170 IFM NANO thrusters have been launched,
in a wide range of different orbits and missions. In-space te-
lemetry of different propulsion systems used has been previ-
ously published –. Figure 20 shows a sample telemetry
of periodical thrusting commanded to 200 µN, and the thrust
achieved. This thrust achieved is calculated by an internal
model based on measuring emitter and extractor voltages and
emitter current. To avoid spacecraft potential drift, the built-
in neutralizer was activated correspondingly. The periodic
thrusting cycles shown are part of a longer firing campaign
spanning over 400 hours, shown in Figure 21 , .
Figure 20. Sample thrusting sequence .
Figure 21. Long-term thrusting sequence .
The large number of thrusters and wide variety of types of
missions and orbits allows to derive lessons learned to im-
prove operational procedures as well as inform future de-
signs. This includes differences in anticipated and actual use
cases, e.g., performing multiple propellant solidification cy-
cles without thrusting, which can have a negative effect on
thruster stability, control aspects including the robustness of
the propulsion system against incorrect commands, as well as
possible interaction with unqualified environment in lower
orbits. The ability to improve operations from learnings
across missions highlights the importance of flexibility of
command software to allow for updates.
7. DEVELOPMENT ROADMAP FOR SCIENTIFIC
The main aim of the recent developments is to optimize the
FEEP thrusters for long duration science and Earth observa-
tion missions, with the development of reliable building
blocks, which can be clustered based on the required total im-
pulse and thrust range. In particular in view of the mission
NGGM, several design upgrades are being implemented and
will be tested in the upcoming years.
Lifetime and Total Impulse
The lifetime requirement for the NGGM mission is 5.5 years,
with the goal of 7.5 years. The endurance test where emitter
A39 has reached 6 years of cumulative firing time has
demonstrated the robustness of the core element of the
The first design modification consists of the increase in total
impulse available for the mission. The most recent NGGM
mission requirement is 50 kNs. To be compliant, a new mod-
ule design able to accommodate 500 g of propellant is being
developed. This would more than double the amount of pro-
pellant available in the currently most popular FEEP thruster
models (220 g) .
As the requirement in terms of thrust range for the NGGM
mission is between 10 µN and 1 mN, two thruster heads
1990 1995 2000 2005 2010 2015 2020
1989 2001 2009 2014
would be sufficient to provide the required thrust with the fo-
cus electrode. However, it is expected that the thrust required
during the NGGM mission is at the lower end of the thrust
range mentioned above. A strong advantage with respect to
other electric propulsion technologies is that the FEEP thrust-
ers can be fired with a thrust as low as 10 µN or less, and with
a thrust resolution of 1 µN. This is mostly thanks to the fine
control possible through the adjustment of the voltage applied
to the emitter.
Power Consumption and Power-to-Thrust Ratio
Another relevant requirement for the NGGM mission is re-
lated to the power budget, which is limited to 40 W/mN with
max. 10 W standby power. The new thruster design will fea-
ture improved thermal insulation to limit the thermal losses
to a minimum. As indicated in section 4, the integrated focus
will allow to increase the PTR by ca. 30%.
Beam Divergence and Thrust Vector Misalignment
A low thrust vector misalignment is crucial for missions such
as NGGM or commercial satellites where the reaction wheels
could reach saturation during long thruster operation periods.
The development of the focus electrode not only provides a
higher PTR by reducing the cosine losses but also helps to
reduce the vertical and horizontal thrust vector angles signif-
icantly as shown in Figure 9.
This paper gives a summary of the developments on the
FEEP technology which have taken place in the last decades
at FOTEC - from the proof of concept of a LMIS up to the
development and successful in-orbit demonstration of an in-
tegrated FEEP-based propulsion system. Over the years, sev-
eral subcomponents have been analyzed and individually im-
proved, showing a promising future for the technology’s
competitiveness in the anticipated increase of market expec-
tations across several crucial metrics such as power-to-thrust
ratio, thrust throttling, low noise, and proper thrust vector
In parallel to the development and advancement of FOTEC’s
proprietary FEEP technology, efforts were made to develop
dedicated diagnostic systems to characterize and to better un-
derstand FEEP thrusters, such as plasma diagnostic systems
(RPA, DFC), the in-situ mass efficiency test stand and multi-
ple generations of µN thrust balances for direct thrust meas-
Several generic use cases as well as some specific upcoming
missions have been identified, which, together with the re-
sults of development iterations, show a clear roadmap for the
future of the presented technology.
This research is the result of a variety of projects funded by
the European Space Agency (in particular
4000104675/11/NL/PA, 4000116898/ 16/NL/EM,
4000133196/20/NL/MM/ff, 400014308/20/NL/FF/tfd) and
the Austrian Research Promotion Agency (874844).
 F. G. Rüdenauer, “Field emission devices for space ap-
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COSPAR Symposium, Singapore, Apr. 20, 2023.
Bernhard Seifert received a Bache-
lor from the Technical University Vi-
enna in Computer Science, a Master
from the University of Vienna in
Physics and a PhD from the Tech-
nical University Graz in Technical
Sciences. He has been working at
FOTEC since 2010 and is the head
of the department Aerospace Engi-
neering. He has developed micro
propulsion solutions based on PPT and FEEP technology. In
2010 Dr. Seifert founded the company silicon systems dedi-
cated to the development of distributed data acquisition solu-
tions for the industry. From 2012 to 2017 he was leading the
development of the first Austrian CubeSat. He is focusing on
the development of the FEEP technology which shall be used
for future science and Earth observation missions.
Laura Bettiol received a Bachelor
and a Master’s degrees in Aero-
space Engineering as well as a
Ph.D. in Space Sciences, Technolo-
gies and Measurements from the
university of Padua, Italy. She has
been working at FOTEC since 2019.
She started as senior scientist,
broadening her responsibilities to
project management and team ad-
ministration over time. She has been involved in several pro-
jects where FEEP thrusters are further optimized and tested,
mainly taking care of risk management, lifetime and perfor-
Nembo Buldrini attained the mas-
ter’s degree in aerospace engineer-
ing at the University of Bologna, It-
aly, with a specialization thesis on
MPD propulsion. From 2001 until
2010, he worked at the AIT Austrian
Institute of Technology (formerly
Austrian Research Centers) as re-
search scientist in the propulsion de-
partment. In 2010 the propulsion
group has been transferred to FOTEC, where he is presently
employed. His main activity consists in the design, assembly
and testing of electric propulsion systems, while keeping an
eye on advanced propulsion concepts.
Martin Eizinger graduated with a
BSc in Mechanical Engineering and
an MSc in Aerospace Engineering in
Austria. He worked at ENPULSION
GmbH for two years and joined
FOTEC in 2022 after another two
years in the Electric Propulsion sec-
tion of the European Space Agency.
His responsibilities include the plan-
ning and execution of tests focused
on research and development, primarily in the fields of elec-
tric propulsion and high-energy particle radiation. Addition-
ally, he maintains and develops software used for data acqui-
sition and processing.
David Krejci received the M.S. de-
gree in applied physics and the
Ph.D. degree in mechanical engi-
neering from Vienna University of
Technology in 2008 and 2012 re-
spectively, and the M.A. degree in
political science from University of
Vienna in 2012. Since 2023, he is
the Chief Science Officer at
ENPULSION, following his tenure
as the CTO since the beginning of the company, including the
launch of 150 flight units. Prior to this, he was a Postdoc and
Research Scientist at the MIT Space Propulsion Laboratory
working on the design and characterization of MEMS-based
ionic liquid electrospray thrusters.
José Gonzalez del Amo is the Head
of the Electric Propulsion Section at
ESA since October 2003 and the
ESA Propulsion Laboratory Man-
ager since 2004. He was before
Electric propulsion Engineer in the
Electric Propulsion Section at ESA
from 1991 to 2003. He started his
career in Space, in 1989, at ESA as
graduated trainee in the Power Con-
version Division. He has a Master degrees in Applied Physics
(University Autonoma de Madrid) and in Space Systems En-
gineering (University of Delft). He has been involved in many
research and development activities on Electric Propulsion
supervising the industrial developments and carrying out
testing and research activities at the ESA Propulsion Labor-
atory since 1991. He has been involved in many ESA projects
such as Artemis, Smart-1, GOCE, Bepi Colombo, Lisa-path-
finder, AlphaBus, Neosat, Small GEO, Electra, Galileo Evo-
lution, etc. He has been also the responsible for the roadmap
of the Electric propulsion Technology within the ESA Har-
monisation programme and the roadmap preparation for all
the propulsion systems called Propulsion 2000. He has near
170 papers in the field. He is also the Coordinator of the Eu-
ropean Commission activity on electric Propulsion, EPIC.
Luca Massotti graduated in Aero-
space Engineering from the Politec-
nico di Torino (IT), and in 2004 re-
ceived his Ph.D. in Aerospace Engi-
neering from the Aeronautical and
Space Department of the Politecnico
di Torino. He was visiting re-
searcher at West Virginia University
to study aircraft modelling and Neu-
ral Network controllers.
He joined Thales Alenia Space in Turin as an engineering
consultant for metrology and AOCS. In 2005 he was a Post
Doctoral Researcher at the EOP directorate at ESTEC facil-
ity and after a consultant at the Future Missions & Instru-
ments division, with RHEA System BV, since 2007. Since
February 2022 he is an ESA staff, appointed as system engi-
neer of the MAGIC/NGGM mission, and coordinator of sev-
eral Inter-Agency working groups between ESA and NASA
on gravity topics.
Dr. Massotti is adjunct professor at Politecnico di Torino,
AIAA Associate Fellow, and lecturer at several universities.
He is author and co-author of more than 100 publications.