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SPACE PROPULSION 2022 Testing the NANO AR³ FEEP cubesat electric propulsion system at ESA Propulsion Laboratory



Space propulsion systems undergo thorough ground testing before being deployed in space. We report the results of a functional verification and performance characterisation test campaign of an integrated electric propulsion system for cubesats and microsats with purely electric thrust vectoring capability and no moving parts. Visualisations of the plume data obtained from Faraday cup scans show a clear, corresponding trend of the variation of the inclination and azimuth angles of the thrust vector when these are commanded. The divergence angle computed from plasma diagnostic data is 49°, independently of the achieved inclination of the ion beam.
Testing the NANO AR³ FEEP cubesat electric propulsion system at ESA Propulsion Laboratory
Martin Eizinger (1), José Gonzalez del Amo (2), Luca Bianchi (2), Davina Di Cara (2),
Quirin Koch (3), David Krejci (3), Alexander Reissner (3), Kaarel Repän (3), Tony Schönherr (3)
(1)(2) European Space Agency (ESTEC), Noordwijk, The Netherlands, Email:
(3) ENPULSION GmbH, Wiener Neustadt, Austria, Email:
KEYWORDS: field emission electric propulsion,
thrust vectoring, characterisation, indium, Faraday
cup, ESA R&D, TDE
Space propulsion systems undergo thorough
ground testing before being deployed in space. We
report the results of a functional verification and
performance characterisation test campaign of an
integrated electric propulsion system for cubesats
and microsats with purely electric thrust vectoring
capability and no moving parts.
Visualisations of the plume data obtained from
Faraday cup scans show a clear, corresponding
trend of the variation of the inclination and azimuth
angles of the thrust vector when these are
The divergence angle computed from plasma
diagnostic data is 49°, independently of the
achieved inclination of the ion beam.
This paper is based on an experimental verification
of the thrust vectoring capability of the DUT.
1.1. Thrust vectoring
Adjustability of the thrust vector of a space
propulsion system is a highly valuable feature.
Depending on the configuration of the propulsion
system on the spacecraft, it allows not only for more
advanced orbital manoeuvres but also for
introducing a torque. Furthermore, a propulsion
system can develop an undesired inclination of the
thrust vector over its lifetime or even present one at
beginning of life (BOL) despite strict manufacturing
tolerances, depending on the technology, e.g. [1],
[2]. This can be compensated only with the ability to
adjust the thrust vector.
Many approaches for thrust vectoring can be
identified. A very low-resolution example is the use
of multiple, spatially separated and selectively
activated thrusters. An asymmetry of the exhaust
plumes naturally results in an off-centred thrust [2].
If the bases of the thrusters are not coplanar, the
vector sum is also inclined with respect to the
spacecraft coordinate system.
A different approach is to mechanically change the
orientation of one or more thrusters, which naturally
brings about an inclined thrust vector. Due to the
complexity of bearings compliant with the
environment of outer space, intricate designs using
compliant mechanisms have been presented that
provide rigid rotation of nozzles around two axes [3].
Another possibility to achieve an inclined thrust
vector is by injecting a secondary flow that diverts
the main flow away from the geometric axis [4].
In contrast to the exhaust gas of chemical
propulsion systems, electric propulsion (EP)
systems have an additional mechanism for applying
forces or torques to the propellant, i.e. the electric
charge of the exhaust particles. In fact, that is how
most EP technologies produce thrust in the first
place, however commonly the resulting force of the
electric or magnetic field is collinear with the
geometrical axis of the thruster.
The device under test (DUT) for which data are
presented in this paper is categorised as an
electrostatic propulsion system. As such, the thrust
it produces is the reaction to the electrostatic force
applied to the ionised atoms in the exhaust plume.
By modifying the electric field to not be
axisymmetric, the ions experience an acceleration
component in radial direction, resulting in an overall
inclined thrust vector.
1.2. Device under test
The NANO AR³ is a fully integrated propulsion
system developed by ENPULSION in cooperation
with FOTEC in the frame of the ESA Technology
Development Element (TDE) project Innovative
Propulsion systems for Cubesats and Microsats. It
is a derivative of the flight-proven Indium FEEP
Multiemitter (IFM) NANO (now commercially called
ENPULSION NANO), which was first deployed in
space in 2018 [5][6] after over 20 years of
collaborative development between FOTEC and the
European Space Agency (ESA) [7][8][9][10][11].
Since then, several new variants of the product
have been in development, including the NANO
AR³, whose primary attribute is its thrust vectoring
The NANO AR³, the device under test, is an indium
Field-Emission Electric Propulsion (FEEP) system
with a crown emitter and a segmented extractor, two
neutralisers and a Power Processing Unit (PPU).
The system provides controllable thrust between
100 and 350 µN, at a specific impulse greater than
2000 s and with a power consumption lower than
45 W. The underlying electric parameters are
entirely controlled by the on-board embedded
firmware through calibrated algorithms that take the
target angles as input.
1.2.1. Basic functionality
During operation and stand-by, the propellant is in
the liquid phase. Capillary forces transport it from
the reservoir to the expulsion area, which is a
circular configuration of 28 small needles, usually
referred to as the crown. The crown is part of the
emitter-subassembly, and entirely wetted with the
propellant, which is indium. It is furthermore
surrounded by the extractor ring, which in case of
the AR³ is split into three segments. Fig.1 shows this
arrangement from an external view of the final
Figure 1: Close-up of emitter and extractor configuration of
the NANO AR³
During nominal operation, the extractors are at a
high negative potential, while the emitter is at a high
positive potential. The resulting electric field applies
a stress to the liquid, electrically conductive
propellant, which consequently deforms at the
needle tips into a shape known as a Taylor cone
[12]. At sufficiently high voltages, a jet of liquid starts
to emanate from the apex of the Taylor cone [12],
and the ions within it are accelerated by the electric
field. Varying the potential across the extractor
segments causes an asymmetric electric field and
consequently a radial component of acceleration.
Independent control of the extractor segments
requires additional electronics compared to the
model with a single extractor.
1.2.2. Propellant
Unlike the most common propellants, indium is not
a gas at room temperature. This, along with some
other differences to noble gases, makes it
convenient in many ways for use as a propellant.
However, it also leads to effects that need to be
considered in testing. For example, Mühlich et al.
developed an advanced design of Faraday cups for
ion current measurement, which is highly accurate
and specifically suitable for indium FEEP ion
sources [13].
1.3. Test objective
The objective of the test is to characterise the thrust
vectoring capability of the DUT. This is achieved by
commanding several combinations of inclination
and azimuth and qualitatively comparing
visualisations of the Faraday scans of these
operating points (OPs). The combinations of
inclination and azimuth of all inclined thrust vector
operating points are listed in Tab.1. The thrust value
varies based on inclination, namely 349 µN, 299 µN,
and 249 µN for 5°, 10°, and 12.5° respectively.
Table 1: Inclination and azimuth for all operating points
The main methods for obtaining test data are
plasma diagnostics, telemetry from the DUT, and
basic data from ground support equipment (GSE).
2.1. Test facility
The test is carried out in the SPF vacuum facility at
the ESA Propulsion Laboratory (EPL) of the
European Space Research and Technology Centre
(ESTEC). The main vacuum chamber has a
diameter of 2 m and a length of 2 m. High vacuum
is mainly achieved with a turbomolecular pump. The
cryogenic pumping systems of the facility are not
used, because of the propellant’s tendency of
sticking to surfaces already at room temperature.
2.2. Plasma diagnostics
In this test, all plasma data are obtained using
Faraday cups (FCs). Measurements are acquired at
multiple elevation angles (up and down relative to
the chamber axis) and sweep angles (left and right
relative to the chamber axis) to achieve a high
resolution of the ion beam. This is realised by a
semi-circular arm that holds FCs and can rotate
around a vertical axis (schematic see Fig.2).
Figure 2: Exemplary scheme of SPF’s diagnostic arm with
Faraday cups
In total, twelve FCs are used to resolve the elevation
from -72° to 72°, while the sweep angle reaches
from -70° to 70° with approximately resolution.
The FCs are distributed asymmetrically to allow for
placement of one probe at 0° despite the use of an
even number of probes. The position of all probes is
summarised in Tab.2.
Table 2: Angular position of Faraday probes with respect to
the horizontal plane
Two different implementations of FCs are used.
More specifically, all but one probe are specifically
designed for use with indium (FOTEC probes). The
twelfth probe’s design is based on noble gases as
propellants (ALTA/SITAEL probe). This is to gain
information about the validity of the probe designs.
Comparability is enabled by placing the ALTA probe
at a location symmetric to one of the FOTEC probes
(specifically 12° and -12°, respectively). Fig.3 shows
one specimen of each probe.
Figure 3: Faraday probes (left: ALTA, right: FOTEC)
The FCs have a collector and a shield, biased at
-20 V, and +10 V respectively.
Data from these probes are acquired in sets called
“scans”, where one scan is one sweep of the arm.
Each scan is stored in a dedicated log file.
2.3. Telemetry
The DUT communicates with the electronic GSE
(EGSE) and transmits a wide range of data such as
housekeeping (e.g. bus voltage and current,
software fuses, temperature of critical components)
and operational quantities (e.g. internal voltages,
thrust derived from a mathematical model, thrust
vector angles).
The software on the EGSE computer makes these
data graphically available to the user and also
stores them in one continuously acquired log file,
split over time for different test segments.
2.4. Ground support equipment
Apart from the DUT-dedicated software, the EGSE
computer runs a software to acquire GSE-related
data and to supply the DUT with power. The DUT is
connected to a laboratory power supply that
simulates the spacecraft bus. In addition, the variant
tested in this campaign has an integrated relay that
serves as an on-off switch, allowing for switch-off of
the device without disabling the bus voltage. A
separate power supply provides the voltage for this
The GSE also measures the temperature at multiple
locations outside of the DUT, including the
temperature reference point (TRP) used for
simulation-based thermal analyses. In addition, the
pressure inside the facility measured with vacuum
gauges is recorded by the software. Similarly to the
telemetry, these data are acquired continuously, but
unlike the DUT telemetry, the data are not split into
test-specific segments.
Diagnostic data are processed in multiple steps
before analysis and presentation. Because these
steps are carried out independently by two parties
(ESA and ENPULSION), they are described
generically rather than with the explicit or
computational operations.
3.1. Axial offset correction
The location on the DUT from which ions are
extracted does not coincide with the centre of the
semi-circular arrangement of probes. Instead, an
offset in axial direction is present as a consequence
of facility-related technical limitations. This results in
multiple geometrical effects that need to be
corrected for.
3.1.1. Different distances of the probes
The effective distance from the location of ion
expulsion to each probe depends on both the
elevation angle of the probe and the sweep angle.
The offset can simply be added to the radius of the
arm if elevation and sweep angle are both equal to
zero (i.e. central probe directly in front of the DUT).
For all other sweep angles, a triangular summation
applies. This triangle only lies in the plane of the
sweep angle for the central probe; for all other
probes, both angles must be considered (see Fig.4).
Figure 4: Effect of a displacement “d” on the perceived
distances and angles of a probe
The distance between the diagnostic arm structure
and the probe collimators is assumed constant for
all probes. This approach neglects the slight
variation resulting from the non-radial orientation of
non-central probes because they are aimed at the
offset point. The final value of the distance is
obtained by taking the absolute of the position
vector of the probe after adding the displacement to
the z component as in Eq.1.
The components are calculated from the arm
radius, its position, and the mounting position of the
probe (see Fig.4):
3.1.2. Narrowing of angles
Both elevation and sweep angle are initially
measured with respect to the centre of the arm. To
convert them to the angles as seen from the location
of ion expulsion, the angles and  of the triangles
described in the previous paragraph are applicable.
Their computation follows from Eq.2, Eq.3, and Eq.4
using simple trigonometry.
3.1.3. Shadowing of the collimator
Another effect of the aforementioned displacement
is that the collimator of the probes appears as an
ellipse rather than as a circle from the point of view
of the location of expulsion. The vertical component
of this effect, i.e. the shadowing pertaining to the
elevation of the probes, is addressed by performing
a laser alignment procedure prior to acquisition.
However, because the offset also applies to the
rotational axis of the arm, this distortion of the
collimator also occurs for the sweep angle. This
results in a reduced intake area of the probe.
Knowing the area perpendicular to the direction of
motion of the ions is critical for correctly computing
the ion current density. Fig.5 shows three examples
of such a distortion (note that the vectors x, y, and z
are to be treated separately for each scenario, i.e.
even though z3 and z2 are drawn the same, they
would not have the same length for position 2 and 3
of any given probe).
Figure 5: Effect of a displacement “d” on the apparent shape
of the collimator (blue) illustrated on the example of three
different probe locations (1: central probe at 0°, 2: elevated
probe at 0°, 3: elevated probe at arbitrary nonzero angle)
Shadowing effects due to the thickness of the
collimator disc are neglected in this analysis. The
area relevant for the calculation of the current
density is the area of the ellipse, which is calculated
as , where is simply the radius of the
collimator and is equal to the radius shortened by
the cosine of the angle between the normal vector
of the collimator (p) and the position vector from the
expulsion point to the probe (r’). Because the
pointing of all probes to the offset point is performed
when the arm is at the centre, the normal vector p
can be computed from the ideal position vector r of
the probe (which results directly from the arm
radius, sweep angle, and elevation angle) by adding
the x and z components of the length d rotated by
the sweep angle. Finally taking the scalar product of
these two vectors yields the factor by which is
shorter than the radius. The area of the visible
ellipse is thus calculated as
 󰆒
with the pointing vector according to Eq.6 and the
position vector according to Eq.1.
3.2. Scaled current density
The aforementioned correction terms as well as
scaling based on distance and area are applied to
the raw currents according to Eq.7, eventually
yielding the current density scaled to a certain
distance from the DUT.
 󰆒
Note that the indexes DUT and “raw” imply the
transformation of angles as described.
This computation assumes that the trajectory of ions
from the point of expulsion to the FC collimator is a
straight line.
3.3. Averaging
Multiple Faraday scans over the whole range of
sweep and elevation angles are performed for each
operating point (OP), where an OP is characterised
by a unique combination of thrust magnitude, thrust
vector inclination, and thrust vector azimuth. The
data of these scans is averaged into one matrix per
operating point.
3.4. Interpolation
The matrix of scaled current densities for a single
OP is fed to a Clough-Tocher 2D interpolation
algorithm [14], which facilitates smoothly plotting the
data as well as integrating in a spherical coordinate
system. The domain of the resulting 2D field is given
by the extrema of the corrected sweep and probe
angles, i.e. two pairs of values.
3.5. Conversion to DUT coordinate system
From the point of view of the DUT, a spherical
coordinate system of zenith and azimuth is more
intuitive than a system of elevation and sweep
angle. This is particularly relevant for the
computation of the divergence angle. Fig.6 shows
the coordinate system into which the data are
Figure 6: Spherical coordinate system centred at the
expulsion point of the device under test and aligned with the
device's Cartesian coordinate system
If the interpolated field is described using this
coordinate system, the definition of its domain
changes slightly: the two pairs of limits become, for
simplicity, one single limit describing the maximum
zenith angle where the field is defined at all
azimuthal angles. This maximum zenith angle 
is the minimum of the absolute values of the four
limits of the original interpolated field.
The aforementioned coordinate transformation is
not performed for the presentation of the raw FC
measurements; instead, these are plotted as the
raw current over the (corrected) sweep angle.
3.6. Divergence angle
The divergence angle is computed based on the
total current and consequently relies on a
summation over the interpolation field. This is done
in spherical coordinates, where the limits of
integration are ideally such that a hemisphere is
covered. However, the zenith angle is limited to
approximately 60° as a result of the range of the
interpolation field.
The mathematical formulation for the integral of the
current density is shown in Eq.8, where the product
of the two terms in parentheses describes the area
of a surface element.
However, because the integration is carried out
numerically, discrete steps of  and  are taken,
causing a small error because the upper edge of the
surface element is shorter than the lower edge. To
minimise the impact of this error, the values for the
current density are taken at the centre of each
discrete step rather than at its edge. The first value
is therefore not taken at like the integral
indicates, but at 
, where  is the step size.
This integral is taken once for  to
determine the total current (at least within the
domain of the data), and then once again with a
break-condition when 95% of the total current is
reached. The angle where this condition is reached
is taken as the divergence angle, equal to the half-
angle of the spherical sector whose cap accounts
for (at least) 95% of the total ion current.
3.7. Measurement offset correction
Performing a scan with the acquisition system on
but with the DUT off (i.e. with no plasma present)
shows a highly stable nonzero signal (see Fig.7).
Figure 7: Current measurements without plasma
This is likely a product of the measurement
electronics, caused by an input offset voltage or an
input bias current through the shunt. If this offset is
assumed to be independent of the amplitude of the
signal, it can simply be subtracted from the
measurement. Fig.8 shows data from one scan,
with and without the offset subtracted. More
specifically, the signals shown in Fig.7 are averaged
over the sweep angle, and these averages are
subtracted from the corresponding channel data.
Figure 8: Current density over sweep angle, plotted by
probe angle, for all probes, with idle offset (left) and with
offsets subtracted (right)
Previous Faraday cup measurements of the plume
of the IFM Nano a comparable ion source have
shown that the current density becomes zero
between 60 and 80 degrees (depending on the
operating point), and that it approaches this limit
linearly [15]. Accounting for the aforementioned
offset, both these characteristics can be seen in the
data of this test. Furthermore, the peak current
density is approximately three times higher than in
the aforementioned study, which correlates with the
three times higher emission current (approx. 3 mA
compared to 1 mA in [15]).
The presented results focus on information inferred
from plasma diagnostic data. For colour plots, the
lower limit is chosen slightly above zero, causing all
negative and zero values to appear white, providing
a stronger contrast along the edges.
4.1. Data examination and verification
Inspection of visualised data initially reveals two
artefacts. The first is a trough in the centre, along
the entire sweep. An example of this is given in
Figure 9: Example case showing a trough at zero elevation
Due to the consistency of the location of this
observation across all operating points, in particular
its independence of the azimuth of the thrust vector,
it can be attributed to the measurement setup, most
likely either to the signal conditioning unit or to the
affected probe(s) themselves. Despite the artificial
nature of this observation, the related data are kept
in the presented results as they are.
Additionally, the data of some OPs show distinct
strands of unusually high or low signal compared to
data in the immediate proximity (example see
Figure 10: Example case showing amplified or weakened
strands with a shape reminiscent of the diagnostic arm
From the shape of these strands, it is evident that
they correspond to one data point of sweep angle
measurements. They appear because scans may
differ in their sampling of the sweep angle, which,
when merging and averaging across multiple scans
of the same OP, may cause individual samples to
appear amplified or weakened relative to the rest of
the scan data.
4.2. Probe type comparison
The measurements of the subject probe (ALTA) is
qualitatively compared to a reference probe
(FOTEC) located at the same angle on the other
side of the mid-plane. This is done by plotting the
raw data of these two probes over the sweep angle.
Fig.11 shows that for an OP with an uninclined
commanded thrust vector, the ALTA probe acquires
a lower current than the FOTEC probe at a
comparable location.
Figure 11: Raw current measured across sweep angle for the
ALTA probe (blue) and a FOTEC probe placed on the
opposite side of the beam (orange)
This pattern of a lower amplitude emerges
consistently across OPs with an inclined thrust
vector, as can be seen in Fig.12.
Figure 12: Raw current measured across sweep angle for the
ALTA probe (blue) and a FOTEC probe placed on the
opposite side of the horizontal mid-plane (orange) for a
strongly inclined thrust vector at multiple azimuth angles
measured clockwise around the thruster axis starting at the
left horizontal, from top left to bottom right
Note here that in the first row of images, the beam
is actually inclined towards the subject, yet the
acquired current is lower than that of the reference,
albeit by less than for an uninclined thrust vector.
Meanwhile, the gap between the two increases
when the beam is inclined away from the subject
(ALTA probe).
4.3. Divergence angle
The divergence angle around the thrust vector is
approximately 49° for all OPs. Note however that
due to the limited field of view of the Faraday scans,
the computation of this angle is an underestimation
of the true divergence angle.
In comparison, Mühlich et al. [15] experimentally
found a divergence angle of 63° for the IFM Nano,
which is the precursor of the DUT. In a simulation of
that device, they found a divergence angle of 49°.
4.4. Thrust vectoring
An excerpt of significant OPs is presented in the
following. The colour scale is kept constant across
all figures, allowing for further comparison. However
this leads to noticeably less contrast for the lower
thrust OPs. Fig.13 shows colour plots of the post-
processed current density of OPs that have a thrust
vector inclination of 12.5°, which is the highest
inclination commanded during this in this campaign.
However, these OPs have the lowest commanded
thrust value, which is 249 µN.
Figure 13: Ion current density over x and y as seen from the
thruster, at 249 µN, 12.5° inclination, and various azimuth
angles measured clockwise around the thruster axis starting at
the left horizontal, from top left to bottom right (red dot marks
commanded thrust vector)
Fig.14 shows the OPs that have a slightly lower
thrust vector inclination (10°) while having an
increased thrust (299 µN).
Figure 14: Ion current density over x and y as seen from the
thruster, at 299 µN, 10° inclination, and various azimuth
angles measured clockwise around the thruster axis starting at
the left horizontal, from top left to bottom right (red dot marks
commanded thrust vector)
Fig.15 shows the OPs with the highest thrust
commanded during this test campaign (349 µN).
However, the thrust vector inclination is reduced to
5°, which makes the skewed current density
distribution less evident from the provided graphs.
Figure 15: Ion current density over x and y as seen from the
thruster, at 349 µN, 5° inclination, and various azimuth angles
measured clockwise around the thruster axis starting at the
left horizontal, from top left to bottom right (red dot marks
commanded thrust vector)
The qualitative characterisation of the NANO AR³
FEEP propulsion system was completed
successfully using the plasma diagnostics setup at
the EPL. Specifically, the inclination of the ion beam
without the use of moving parts was achieved and
verified by means of Faraday cup data. The beam
divergence is not measurably affected by the
inclination of the beam, and it is comparable to the
divergence angle obtained during a previous,
independent study on a precursor model.
The ALTA Faraday probe designed for use with
Xenon acquires a lower current than the FOTEC
probes, which are designed for use with indium.
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Field emission electric propulsion thrusters are characterised by their low thrust range, which makes them ideal for the precise control of spacecraft. Decisive for precise control are the properties of the thruster ion beam, which includes the beam divergence angle and the thrust vector. The analysis of these properties is also necessary in order to be able to estimate the interactions of the beam with components of the spacecraft. Due to such interactions, solar panels or electrical instruments on board the spacecraft could be damaged by sputtering effects. The spatial ion current density and energy distribution of a test crown emitter beam, with different specifications compared to the IFM Nano thruster, were examined experimentally with a diagnostic system, including Faraday cups and a retarding potential analyser. In addition to the analysis of the beam profile of an emitting crown, a single emitting needle was analysed. Based on these experimental analyses, an ion trajectory simulation model was developed to determine the theoretical ion current density distribution. This model includes the properties of a liquid metal ion source, where the ion trajectories start from their point of origin, the so-called Taylor cone jet cap. The benchmark of the model shows that the thrust vector and divergence angle correspond to the experimental results and shows the identical calculations for different thruster parameters, like emission current and electrode voltage. The simulation allows for the optimisation of existing and novel thruster geometries in terms of performance, reliability and longevity.
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The IFM Nano Thruster is a variable specific impulse electrostatic thruster based on Field Emission Electric Propulsion (FEEP), in which a liquid propellant is electrostatically extracted and accelerated to high exhaust velocity. The core element of this propulsion technology is a passively fed, porous ion emitter consisting of 28 sharp emitter tips. This emitter technology has been developed and qualified over decades at FOTEC and the Austrian Institute of Technology, and has recently been adapted for the use as main propulsion system in Nano-and Small-satellites. The resulting IFM Nano Thruster occupies approximately 0.8 U and can be operated between 10 and 40W, resulting in thrust of up to 0.35mN. The thruster can be operated at specific impulse levels between 2000s and 6000s, adapting to mission needs as well as power availability, allowing for significant throttling capability between a couple of µN and 0.5mN. Due to the high specific impulse and high propellant density, the thruster can produce total impulses between 5000Ns and beyond 12000Ns when operated at specific impulses at 2000s and 5000s respectively. The first IFM Nano Thruster has been successfully integrated into a commercial 3U CubeSat in 2017 after undergoing environmental testing, and was launched in January 2018 for a first in-orbit demonstration (IOD). This IOD represents the first instance of a liquid metal FEEP thruster to be operated in space as a primary propulsion system. This paper will present the FEEP thruster principle and experimental characterization, including the first in-orbit test results.
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Summary A variety of multi-emission Indium Liquid-Metal-Ion-Sources have been manufactured and investigated. It was demonstrated, that such configurations are possible and allow very high peak thrusts up to 0.5 mN and possibly above. The prototypes also show promise to enable high mass efficiencies at thrust ranges in the tens of µN range, as required by missions such as SMART-2. If successful, also future Power-Processing-Units (PPU) can be simplified by using integrated multiple In-FEEP thrusters instead of clustering of individual In-FEEP thrusters for high thrusts. Acknowledgement 4All past research activities were mostly carried out by M. Fehringer and F. Rudenauer. This work wass supported in part by ESTEC-Contract 12376/97/NL/PA. 5 References 1 Genovese, A., Steiger, W., and Tajmar, M., "Indium FEEP Microthruster: Experimental Characterization in the 1-100 µN Range," AIAA Paper 2001-3788, 2001 2 Tajmar, M., Genovese, A., Steiger, W., "Indium FEEP Microthruster Experimental Characterization," Journal of Propulsion and Power, submitted (2002)
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The Laser Interferometer Space Antenna project (LISA) is a co-operative program between ESA and NASA to detect gravitational waves by measuring distortions in the spacetime fabric. LISA Pathfinder is the precursor mission to LISA designed to validate the core technologies intended for LISA. One of the enabling technologies is the micro-propulsion system necessary to achieve the uniquely stringent propulsion requirements. Two competing systems, a cesium slit emitter (Alta, Italy) and the indium needle emitter (AIT, Austria) technology have been commissioned to develop this micro-propulsion system. At this point, the cesium slit emitter was chosen by ESA as baseline. The indium needle emitter technology was chosen as back-up and its development and test are still proceeding and the obtained results are documented in the present publication. In the framework of the qualification tests for LISA PF, AIT conducted an endurance test over 3650 hours collecting a total of 586 Ns. A flight representative thruster unit (TCA) was used in this test. The test consisted of three phases. In the first phase the thruster's performance was investigated with a standard laboratory power supply and EGSE. In the second phase the thruster was controlled by the flight representative EGSE and PCU provided by Galileo Avionica S.p.A. Finally, in the third phase a long duration assessment of the thruster was conducted. During this phase (roughly 3200 hrs) the thruster was operated in a mode similar to the one foreseen for operation in space. Additionally, frequent performance and health checks were conducted. Compliance to all requirements subject of those three test phases has been shown. The thruster has an exceptionally low rate of sparks, its noise is well below the stringent noise requirements of LISA, the maximum thrust step resolution is roughly 0.1 μN and the capability to cover the complete required thrust range of 0-100μN has been shown as well as peak thrust values of 150μN. An investigation of the potential impact of a thermal cycle showed that the thruster performance does not change when exposed to such a procedure (e.g. spacecraft separation during the LISA PF mission). Based on the performance of the thruster during the 3650 hours, the expected lifetime of the system was extrapolated to 39,000 hours generating a total impulse of 6300 Ns at an average thrust of 50μN.
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This paper describes the conception, modeling, and development of a fully compliant two-degree-of-freedom pointing mechanism for application in spacecraft thruster, antenna, or solar array systems. The design objectives and the advantages of a compliant solution are briefly discussed. Detailed design decisions to meet project objectives are described. Analytical and numerical models are developed and subsequently verified by prototype testing and measurements in several iterations. A final design of the 3-D printed titanium monolithic pointing mechanism is described in detail and its performance is measured.
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The increasing application of microsatellites (from 10 kg up to 100 kg) as well as CubeSats for a rising number of various missions demands the development of miniaturized propulsion systems. Fotec and The University of Applied Sciences at Wiener Neustadt is developing a number of micropropulsion technologies including both electric and chemical thrusters targeting high performance at small scales. Our electric propulsion developments include a series of FEEP (field emission electric propulsion) thrusters, of which the thrust ranges from μN to mN level. The thrusters are highly integrated into clusters of indium liquid-metal-ion sources that can provide ultralow thrust noise and long-term stability. We are also developing a micro PPT thruster that enables pointing capabilities for CubeSats. For chemical thrusters, we are developing novel micromonopropellant thrusters with several hundred mN as well as a 1–3 N bipropellant microrocket engine using green propellants and high specific impulse performance. This paper will give an overview of our micropropulsion developments at Fotec, highlighting performance as well as possible applications.
ESA’s Next Generation Gravity Mission (NGGM) is a candidate Mission of Opportunity for ESA-NASA cooperation in the frame of the MAss change and Geosciences International Constellation (MAGIC) . The mission aims at enabling long-term monitoring of the temporal variations of Earth’s gravity field at relatively high temporal (down to 3 days) and increased spatial resolutions (up to 100 km) at longer time. Such variations carry information about mass change induced by the water cycle and the related mass exchange among atmosphere, oceans, cryosphere, and land, and will complete our picture of global and climate change with otherwise partial or unavailable data. Over the last 15 years, numerous system and technology activities have been initiated by the Earth Observation Programmes (EOP) Directorate of the European Space Agency with the aim of advancing the maturity of the NGGM system and the key subsystems: particular attention was devoted to the design of the fine attitude control system, enabled by proportional thruster like variable specific impulse electrostatic thrusters based on Field Emission Electric Propulsion (FEEP), in which a liquid propellant is electrostatically extracted and accelerated to high exhaust velocity. The core element of this propulsion technology is a passively-fed, porous tungsten crown emitter, consisting of 28 sharp needles. This emitter technology has been developed and qualified over more than a decade at FOTEC and the Austrian Institute of Technology under ten ESA/EOP contracts since 2005 targeting the NGGM needs and it has recently been adapted for use as the main propulsion system in commercial nano- and small-satellites. This paper summarizes the development efforts of the last decade and provides an assessment of the performance of this thruster technology, after extensive simulation and testing campaigns.