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IAC-02-S.2.07
THE MERITS OF COLD GAS MICROPROPULSION
IN STATE-OF-THE-ART SPACE MISSIONS
Hugo Nguyen, Johan Köhler and Lars Stenmark
The Ångström Space Technology Centre, Uppsala University,
Box 534, SE-751 21 Uppsala, Sweden.
Hugo.Nguyen@angstrom.uu.se Fax: +46 18 471 3572
Abstract: Cold gas micropropulsion is a sound
choice for space missions that require extreme
stabilisation, pointing precision or contamination-
free operation. The use of forces in the micronewton
range for spacecraft operations has been identified as
a mission-critical item in several demanding space
systems currently under development.
Cold gas micropropulsion systems share merits with
traditional cold gas systems in being simple in
design, clean, safe, and robust. They do not generate
net charge to the spacecraft, and typically operate on
low-power. The minute size is suitable not only for
inclusion on high-performance nanosatellites but
also for high-demanding future space missions of
larger sizes.
By using differently sized nozzles in parallel
systems the dynamic range of a cold gas
micropropulsion system can be quite wide (e.g. 0 –
10 mN), while the smallest nozzle pair can deliver
thrust of zero to 0.5 or 1 mN using continuously
proportional gas flow control systems.
The leakage is turned into an advantage enabling the
system for continuous drag compensation. In this
manner, the propellant mass efficiency can be many
times as higher than that in a conventional cold gas
propulsion system using ON-OFF-control.
The analysis in this work shows that cold gas
micropropulsion has emerged as a high-performance
propulsion principle for future state-of-the-art space
missions. These systems enable spacecraft with
extreme demands on stability, cleanliness and
precision, without compromising the performance or
scientific return of the mission.
Copyright
Copyright 2002 by Hugo Nguyen. Published by the
American Institute of Aeronautics and Astronautics,
Inc., with permission. Released to IAF/IAA/AIAA
to publish in all forms.
1. INTRODUCTION
For Attitude and Orbit Control System (AOCS)
requirements on extreme stabilisation, pointing
precision, contamination-free operation is the
most important for many missions. Examples
include DARWIN, LISA, SIM, NGST, and those
which have optical or other equipments pointing
exactly to a certain object or in a certain direction.
Other obviously desirable properties include
design simplicity, cleanliness, safety, robustness,
low-power operation, no net charge generation to
the craft, together with low mass, and a wide
dynamic range. Cold gas micropropulsion has
those merits and in the present paper,
technological solutions will be treated
emphatically, at the same time the mentioned
merits will be brought out clearly.
2. SYSTEM DESCRIPTION
The nanosatellite designed at The Angstrom
Space Technology Centre, Uppsala University,
Sweden (ASTC) has four identical thrusters units,
or thrusters pods (figure 1a). Each unit constitutes
an autonomous cold gas micropropulsion system,
which is presently under development. The
spherical housing (figure 1b and c) accommodates
four identical thrusters. The unit also contains
electronics for local closed control loops of thrust
and a serial data interface to the satellite AOCS.
Each thruster is a complete microsystem,
including a nozzle with internal heater, a
proportional flow control valve, particle filter, and
sensors for pressure, temperature and thrust. The
total mass of the thruster unit is below 60 grams
and the diameter of the spherical envelope is 41
mm. All gas control is performed in a central
stack consisting of four silicon wafers. A common
main shut-off valve is also included in this unit.
2
All control electronics are located as hybrids on
three individual wafer stacks underneath the central
stack. Each of these stacks is composed of three
silicon wafers
a
b c
Figure 1: a) The nanosatellite under development at
Angstrom Space Technology Centre, Uppsala
University, Sweden. b) and c) The 60-grams
autonomous thruster unit with four nozzles, designed
for thrust level 0-10 mN.
3. COLD GAS MICROPROPULSION AND THE
MERITS
Cold gas as fuel
Cold gas is, by definition, not warm. The terms
warm and cold on Earth usually refer to room,
human body or water freezing temperature. For cold
gas as propellant of propulsion system the
temperature is not standardized. Nevertheless, two
things are relevant for definition. The propellant has
to be in gas phase when rushing out from thruster,
and no combustion should occur. Yet, thawing and
warming up are not restricted to the propellant to be
called cold gas. Resistor jet propulsion in that
classification will be an exception case.
Higher atom weight of the expelled gas is desirable
due to the third Newton law. Together with the
contamination free constraint (treated briefly below),
the moderate low boiling and melting temperature
are desirable features for a propellant gas from
system point of view, where mass efficient storage
of the gas is a major concern.
Xenon is a potential cold gas propellant, that is, a
heavy and inert gas (table 1). However, there are
some challenges when using this gas, since its
viscosity increases with temperature within a
certain range, that implies that the specific
impulse, I
sp, for a Xenon system becomes very
low if heating the gas.
Helium and Nitrogen are much lighter than Xenon
and their temperatures for storage in liquid or
solid form is technically more demanding, but the
specific impulses are higher.
Propane, C3H8, has been used as cold gas
propellant in microsatellites in past decades.
Butane, C4H10, has been demonstrated
successfully as propellant in the cold gas
propulsion system on SNAP-1, a 6.5 kg-
nanosatellite. However, these gases are organic
compounds and hazard classified.
Carbon dioxide, CO2, has the advantage to
sublimate directly from -78°C at a pressure of 1
bar and the gas has a fairly high specific impulse.
Table 1: Cold gas propellant performances2,3,9
Mr = Molecular weight. tm = Melting temperature. tb = Boiling
temperature.
ρ
= Density (241 bar, 0oC). Isp,t = Theoretical specific
impulse. Isp,m = Measured specific impulse
Cold
gas
Mr
kg/kmol
tm
(1 bar)
oC
tb
(1 bar)
oC
ρ
(241 bar)
g/cm3
Isp,ta)
s
Isp,ma)
s
H2 2.0 -259 -253 0.02 296 272
He 4.0 -272 -269 0.04 179 165
Ne 20.4 -249 -246 0.19 82 75
N2 28 -210 -196 0.28 80 73
Ar 39.9 -189 -186 0.44 57 52
Kr 83.8 -157 -152 1.08 39 37
Xe 131.3 -112 -108 2.74b) 31 28
CCl2F2 121 -158 -29.8 --- 46c) 37
CF4 88 -184 -128 0.96 55 45
CH4 16 -182.5 -161.5 0.19 114 105
NH3 17 -78 -33 Liquid 105 96
N2O 44 -91 -88 --- 67 c) 61
C3H8 41.1 -187.7 -42.1 Liquid --- ---
C4H10 58.1 -138.3 -0.5 Liquid --- ---
CO2 44 --- -78 (S) Liquid 67 61
SF6 146.1 --- -64(S) --- --- ---
a) At 25oC. Assume expansion to zero pressure in the case of the
theoretical value.
b) Likely stored at lower pressure value (138 bar) to maximize
propellant-to-tank weight ratio.
c) At 38oC (560R) and area ratio of 100.
(S) Sublimation
Sulfur hexafluoride, SF6, is one of the most
interesting gases owing to its heavy molecular
weight – heavier than Xenon – and it sublimates
3
at barely -64°C at a pressure of 1 bar. SF6 is non-
flammable and not classified as a toxic gas. SF6 is
believed to be a strong candidate for cold gas
micropropulsion systems under development.
Propellant storage in solid phase, especially
sublimating substances, onboard a spacecraft is
favorable in comparison to storage in liquid phase.
Liquid propellant in ordinary tank is known for
causing sloshing problem, which can severely
disturb high precision stabilization and pointing.
The gas propellant in our system will preferably be
stored in solid or liquid phase. It will be transformed
into gas phase before leaving the storage tank via the
feeding lines, through a micromachined filter set and
the proportional valve, and into the heat exchanger
chamber for warming up. The gas temperature
before leaving the nozzles may reach many hundreds
degree C. Here, the material properties of the silicon
structure at elevated temperatures are the major
limiting factors to the allowed heating. Silicon is a
good thermal conductor. The heat that should go to
the gas can be lost into the wafer stack, seriously
degrading the heater efficiency and possibly causing
secondary heat induced failures in nearby
microstructural elements. Thermal aspect in design
and material selection is therefore an inevitable
matter of concern in the proper engineering of
microsystem.
The system specific impulse, Issp, as a function of
delta-v has been presented by Köhler1. There
different parameters of the system, such as available
storage techniques, dimension of microfluidic
system and nozzles, expelled gas viscosity, pressure
control system, etc., was not taken into consideration
in order to gain the maximum system specific
impulse. Yet, the analysis showed how to estimate a
system specific impulse versus delta-v using cold
gas micropropulsion system in feasibility study for a
certain mission.
Simple design, clean, safe, robust, and low cost
Conventional cold gas propulsion systems have been
used successfully since many years, for example on
Astro-Spas, Hipparcos, EURECA, CHAMP and
GRACE. Cold gas propulsion system is just a
system that controls pressurized gas through a
number of nozzles. It is the simplest form of rocket
engine. By using inert gas it represents one of the
cleanest and safest systems. However,
conventional cold gas propulsion systems have for
many years been considered “old fashioned”,
mainly because of the low efficiency, which
makes the propellant gas supply bulky and heavy.
Also, due to the valve technology traditionally
involved, a minimum impulse bit obtained is
disturbingly high in many applications.
Microelectromechanical system (MEMS, or MST
– Microsystem technology, which is another
abbreviation for the same technology) provides
the possibility to build a high performance system
for use in many new demanding applications1,10.
Cold gas propulsion systems are preferable in
many cases where cleanliness, simplicity, and
reliability are more important than other qualities.
Due to their simplicity the conventional cold gas
propulsion systems have even the lowest cost of
manufacturing. However, this does not hold true
in the same proportion when employing MEMS-
technology for manufacturing.
The desirable gas leakage – Proportional Gas
Flow Control (PGFC)
Leakage always occurs when gas pressure inside
and outside of a system differs. The propulsion
system containing gas of high pressure in a high
vacuum like space will be subjected to
considerable leakage problems, mainly at its
valves and connectors. For conventional
spacecraft the acceptable leak rate has been found
at 10-3–10-4 scc/s GHe (gaseous Helium)4. For
nanosatellites the leak rate is often calculated
from the thrust generated just by the leak and it
should be set to a lower value than the limit
mentioned above. Of course the sum of leaked
mass of gas has to be compared with the amount
of gas on board for the whole mission duration.
Numerous variants of MEMS-based valves have
been developed over the years, with varied
solutions for actuation. The review by Mueller
presents many of them5. The main principles of
valve seat and lid features are also discussed here.
The interesting approaches in order to suppress
gas leakage are making the valve lid from a hard
material that presses against a softer seat, or else a
hard valve lid with a sharp circular edge that
presses against a hard seat, crushing particles that
obstruct the closing of the valve. Both valve lid
4
and seat surfaces can also be made perfectly flat
from hard material. The risk for spontaneous
bonding is obvious here; hence different thin film
coatings have to be employed. For all of the
mentioned approaches the fatigue and wear are the
other problem of great concern.
In our system approach, the drawback of gas leakage
has been turned into an advantage by controlling the
leak rate. Two different operation modes have been
designed for the valves system. The first operation
mode is called Proportional Gas Flow Control
(PGFC), in which the valves will not be fully closed.
Since the pairs of MEMS-based opposite thrusters
are placed symmetrically in the same unit, leakage
from them can be balanced. The valve actuators
have been designed to operate at 10 Hz with
virtually continuous stroke control. The PGFC mode
works on a fairly low electrical bias and actually
consumes negligible electric power (discussed
below). The extremely low gas flow used in this
mode make the continuous operation acceptable in
terms of propellant mass efficiency.
The second operation mode is the shut down mode,
which needs higher electric potential to assure the
complete closing of the valves. Since this mode is
only rarely used, when for instance drifting freely or
if rebooting of satellite is required. In this way, the
fatigue and wear problem is reduced almost to zero.
Spacecraft, especially those on Low Earth Orbit
(LEO), experience a certain drag, which implies that
their attitude has to be corrected. Conventional
AOCS with cold gas propellant operate by the ON-
OFF principle, which is known as a bang-bang-
system. Mission requirement on pointing precision
determines the minimum impulse bit of the AOCS.
For an ON-OFF operating system the propellant
used during minimum impulse bit generation is not
very efficient due to the response delay of valve
actuation, and gas flow. The cold gas
micropropulsion system developed at ASTC with
continuous PGFC meets the ideal requirement on
pointing precision, that is, continuous drag
compensation. In other words, the minimum impulse
bit does not exist. The minimum gas amount does
not leak undesired through the valve system, but is
meant to be released variably in order to counteract
the disturbances continuously. The use of gas
propellant in this manner is estimated three or four
times as effective as that in an ON-OFF system. It is
in fact legitimate to maintain that the gas leakage is
desirable. For a cold gas propulsion system with
continuous PGFC the common opinion about the
propellant inefficiency of a cold gas propulsion
system does not hold true.
Extreme stabilisation and pointing precision
Cold gas micropropulsion system that operates in
the continuous PGFC-mode eliminates the
minimum impulse bit. The resolution of a single
unit, containing two pair of opposite thrusters, is
only dependent on the control voltage applied to
the valve actuators. This technological feature
makes the extreme stabilisation and pointing
precision of a satellite possible. Furthermore, the
broad dynamic range and particularly quick
transition from one thrust level to another –
simply by changing the valves actuation voltage –
make this cold gas micropropulsion system
unique regarding the precision control of the
satellite.
Nevertheless, realisation of a microfluidic system
encounters some delicate problems due to fluidic
mechanical behaviours, such as pulsation and
turbulence in larger system sections, which
constitute the system noise. This in turn
determines the lowest flow rate of a single
thruster, and that implies the minimum continuous
consumption of propellant.
Many experiments and measurements in space
demand vibration-free environment for the
payload and pointing precision of spacecraft.
When free drift does not meet the requirements
the spacecraft has to be actively and continuously
controlled by thrusters. The thrust levels for mini
and microsatellites in that case must be very low
and the thrust resolution has to be below 1 mN.
Gravity and Ocean Circulation Explorer (GOCE),
for example, is a minisatellite of 800 kg under
development. It will operate in LEO (250-300
km) and will be subjected to residual air drag.
Two ion thrusters with a thrust between 1-12 mN
will be used for the drag compensation. However,
small variations of both thrust and thrust vector
call for an additional three-axis system with 0.7
mN maximum thrust and a resolution of 0.25 µN
(see more in references 6 and 7). A cold gas
micropropulsion system investigated at ASTC
would be one of the suitable AOCS for the
spacecraft, since it has a 12-bits wide dynamic
5
range. By employing differently sized nozzles in
parallel systems the dynamic range of the cold gas
micropropulsion system can be increased to 0-10
mN, where the smallest nozzle pair can deliver
thrust of zero to 0.5 or 1 mN using PGFC. A single
smallest nozzle works with a resolution of 0.2 µN.
The influence of mass in motion when actuating the
valve lids on spacecraft stabilisation and pointing
precision was a subject of discussion. However,
calculation for piezoelectric actuated valve designed
at ASTC (figure 2) showed that the valve will
contribute an acceleration of <10-22 g, while the
requirement on maximum vibration level for GOCE,
for instance, was set to <10-15 g. Therefore, redesign
of the valve in order to balance its mass in motion
was not considered necessary.
Figure 2: Piezoelectric proportional valve (14x5
mm) with microfabricated silicon parts
Contamination-free operation
The expelled particles from firing thrusters for
attitude control or when leaking in unfavourable
condition, for instance while the satellite is in
shadow and has no surrounding plasma, can follow
the satellite for a certain time. Their chemical
reactivity towards materials on satellite surfaces can
play a crucial roll in propellant selection. If an
expelled particle is statically charged, it can be
attracted to the satellite surface. This matter raises
concerns due to the size of nanosatellites and its
surfaces of multifunctional films and devices. A
single defected spot on the very limited
multifunctional surface of the tiny satellite is more
harmful than on the larger surface of a conventional
satellite. Contamination problems due to propellant
type are also severe in cases where spacecraft carry
optical instruments with sensitive surface of lenses
and mirrors. Furthermore, if spacecraft in close
formation within distances of 25 to 50 metres, for
instance, the risk of surface contamination on
neighbours in the constellation is not negligible.
No net charge generation to the spacecraft
The cold gas micropropulsion system does not
ionize the expelled gas particles, thus efforts and
power needed for active plume neutralisation can
be spared. A charged spacecraft in relation to
space particles and the expelled gas tends to
attract them to its surface as mentioned above.
In despite of the fact that the spacecraft will be
subjected to electromagnetic forces that would
cause incorrect pointing, these forces are too
small to significantly increase the consumption of
propellant for attitude control. A rough calculation
of the electromagnetic force F that exerts on a
charged spacecraft in LEO can be made as follow.
Assume that the spacecraft is spherical with radius
r=1m, moving with velocity of v=10 km/s in LEO
with magnetic flux density B=5·10-5 T. The
spacecraft is charged to U=1 kV. The maximum
value of the force will be F = Q·V·B = C·U·V·B =
4
πε
or·U·V·B = 5.56·10-8 N = 5.56·10-2 µN. Further
calculation for 10 year mission assuming that the
Isp= 700 Ns/kg for N2 in mission average, shows
that an additional 0.025 kg of propellant would be
needed to counteract the mentioned
electromagnetic force. The same calculation for
monopropellant Hydrazine, N2H4, with Isp=2150
Ns/kg results in 0.008 kg increase of the
propellant. This rough estimation shows that the
force does not constitute any problem in
propellant budget for attitude control. However, it
is comfortable to realize that by using a cold gas
we do not need to care about the impact of
electromagnetic force on attitude control, while
the propulsion system can be kept simpler in
design.
Low-power operation
The electric power needed for operation of a cold
gas micropropulsion system is mainly determined
by the valve actuators and the electronic system.
In the ASTC design, the piezoelectric elements in
the valve actuator work at 0-50 volts. They
function like a capacitor, that is, they consume no
power in the charged state at constant voltage.
The capacitance of a single piezoelectric element
6
is 130-250 nF, depending on size and operation
temperature. Thus, the actuation power is very low.
Changing from one thrust level to another does not
require any considerable consumption of power.
In comparison to conventional satellites and other
spacecraft a nanosatllite should have a limited
number of AOCS in order to minimize the total
mass. Momentum flywheel system for attitude
correction can be omitted to make the satellite more
robust, and to save mass and power consumption.
The cold gas micropropulsion system itself, with its
large dynamic range, covering thrust level from zero
to 10 mN, already guaranties the required precision
of the attitude and orbit control. Nevertheless, a
simple magnetorquers system, that does not require
much power, space, or mass on board, can be
justified as a complementary system in order to save
a significant part of available propellant.
Silicon microsystems can be much smaller
Industrial standard NC-machines have tolerance
0.01 mm or slightly narrower. Today, there are
machines that goes with 75,000 rpm or twice that
speed, and they are capable to mill or drill a row of
holes in a human hair without missing it. Yet, as the
dimensions of work pieces shrink in order to
miniaturize a device, more exotic machining
methods have to be employed, such as laser or
electron beam machining. However, those
machining methods, especially when material is
removed mechanically, leave residual stress and
coarse surfaces. These flaws are not negligible in the
small features concerned.
Silicon, with its perfect atom lattice, offers
incredible mechanical, electrical, chemical, and
thermal properties compared to conventional
mechanic construction materials. The main
manufacturing methods are different etching,
deposition, and bonding techniques. By choosing
proper methods, chemicals, and combinations of
materials the residual stress and other problems can
be avoided or minimized. Silicon microsystems
technology is not restricted to fit standard
component or tools, but rather to the perfection of
the atom lattice and the precision of pattern transfer.
Dimensions can be reduced from few millimetres
down to size of few atom layers.
Mass reduction
One of the most important advantages in silicon
microsystem technology is the possibility to
integrate a multitude of MEMS-devices and
electronic chips. Here, it is not the matter of
miniaturization, but “microturization”. MEMS
technology offers many advantages, such as
higher level of integration of devices, higher
sensitivity and reliability, better mechanical
properties of materials, and many other properties,
which do not show significantly or not at all in
macro-mechanical designs. Of course there are a
number of aspects that have to be taken into
account when using MEMS-solutions. For
instance, flat surfaces of thin element can be
bonded spontaneously; toxic gas, substances, and
solutions must be employed in manufacturing.
Furthermore, the yield of the manufacturing
process typically decreases with increasing level
of system integration.
To illustrate the possible level of device
integration we can consider how many mass
percent is the “payload” in a Pentium processor,
that is, the total mass minus the mass of the carrier
substrate and the plastic envelope. This is
compared with the ratio in case the processor is
built of individual transistors, capacitors, resistor,
circuit board, and solder. The comparison can be
made in the same manner to grasp the mass saving
possible by the transition from conventional
mechanics to microsystem integration of MEMS
devices.
The autonomous cold gas thruster unit designed at
ASTC for AOSC weights lees than 60 grams
(figure 1b and c), of which 40 grams is the
AA7075-aluminum housing. The aluminium mass
in the present design is mainly needed for
radiation protection reason. Both crumb and
housing are still targets for further mass reduction.
A bulkier hybrid unit has previously been
designed without built-in control electronics, and
weighs around 150 grams (figure 3).
The mass reduction naturally saves launch cost,
but can also be used to incorporate more
redundancy by allowing parallel systems. For
instance, in a satellite swarm, loss of one or few
satellites would not jeopardise the whole mission.
On individual spacecraft, the vital microsystem
may also be multiplied without incurring mass
7
penalty. Function failure at a thruster or even a
thruster unit would not erase the whole operation of
the spacecraft.
Figure 3: 150-grams hybrid thruster unit without
build-in control electronics, designed for thrust level
0-1 mN or higher
The launch cost per mission using nanosatellites and
State-of-the-art launch vehicles is low, but the cost
per kilo of go-to-orbit mass is much higher
compared to using a huge lifting rocket for a
satellite of hundreds or thousands times more than
the nanosatellites mass8. However, there is a belief
that strong and fast development of nanosatellites
will initiate an evolution of new launching
technology that suits tiny satellites. Launching such
satellites may not have to be from a launch pad. It
can be made from a jetfighter or from something
else that we just cannot imagine today.
Examples on future demanding space missions
Using components off-the-shelf is the most practical
approach to build a spacecraft, or anything else, due
to well known performances of the components. The
time-to-completion and costs for a planned mission
can reduce substantially. However, when the
components do not meet the mission requirements
new components or the whole system have chance to
emerge. New technological challenges will be
overcome by creative ideas. New missions require
such technologically demanding spacecraft, which
can help us to understand the solar and interstellar
systems. Examples on demanding future missions
are LISA, DARWIN, SIM, NGST and many other
planned and proposed missions. LISA (Laser
Interferometer Space Antenna) has the task to
detect gravitational waves from massive black
holes and galactic binaries. DARWIN aims to
look for life. Both missions will use flotilla of
spacecraft, flying in formation in vast distances of
thousands kilometers from each other with an
accuracy of less than 20 microns. The SIM (Space
Interferometry Mission) will carry a telescope
which can determine positions and distances to
stars with an accuracy several hundreds times
greater than current telescope technology allows,
in order to street-map our Milky Way galaxy. The
NGST (Next Generation Space Telescope) will
look into our history back to the Big Bang with a
telescope that can capture images beyond the
visible portion of the electromagnetic spectrum.
LISA and DARWIN are conducted by European
Space Agency (ESA), while the SIM and NGST
are conducted by the American NASA.
Considering that the attitude control subsystems
currently available would provide 15 arc-seconds
pointing precision per axis, the AOCS with
continuous thrust control, using cold gas
micropropulsion, should be a tempting offer.
Improvement of cold gas micropropulsion system
Up to the present time, the MEMS technology that
produces autonomous ad hoc systems with several
micromechanic and electronic parts, highly
integrated into a single chip, is not matured,
neither in fabrication nor in attitude of the
presumptive users. Cold gas micropropulsion
systems are not any exceptions. However, the
traditional flaws and weakness of the system will
eventually be overcome. For instance, heating of
propellant gas is limited by silicon structure. At
temperature 5000C and above the mechanical
properties of silicon will be degraded. The
measure for this problem may be replacement of
silicon by another material. The heat loss into the
structure when heating the propellant gas is a
reason for a new creative design, in order to make
the specific impulse of cold gas higher11,12. The
8
fabrication yield will be significantly improved as
the manufacturing steps and material systems are
optimized. However, the progress in research and
development support the strong belief that cold gas
micropropulsion systems will be flight-proven
within this decade.
4. CONCLUSION
The cold gas micropropulsion system benefits
greatly from using a continuously proportional
control on the thrust. The configuration of two
opposite thrusters in the same unit balances the
thrusts against each other, so the impulse obtained as
the difference of them may be reduced to zero.
Simultaneously, any troubles emerging from
extremely low flows at near-zero thrust from a
single thruster can be avoided. The same strategy
can be extended to continuous operation, never
closing the valves completely. In this way, leakage-
induced thrust fluctuations are bypassed, and
sticking problem and wear of the valve seats is
heavily reduced. The extremely low gas flow used in
the system make this continuous operation
acceptable in terms of propellant mass efficiency.
Main concerns are the low specific impulse and the
consequent large amount of propellant required.
System analysis shows that low delta-v missions or
their equivalent (e.g. attitude and orbit control) are
suitable missions for cold gas micropropulsion. The
choice of propellant is naturally of prime importance
– carbon dioxide and sulfur hexafluoride is possible
to store in solid form to avoid sloshing problem,
while liquid gas may act as a cooling fluid for the
scientific payload in addition to being used as
propellant.
The analysis in this work shows that cold gas
micropropulsion has emerged as a high-performance
propulsion principle for future state-of-the-art space
missions. These systems enable missions with
extreme demands on stability, cleanliness, and
precision, without compromising the performance or
scientific return of the mission.
References
1 Köhler, J., et al. A Hybrid Cold Gas Micothruster
System for Spacecraft. Sensor and Actuators. A97-
98, pp 587-598, 2002
2 Ayward G. And Findlay T. SI Chemical Data -
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3 Mueller J., in Micropropulsion for Small
Spacecraft, Ed. Michael M. et. al., Chapter 3:
Thruster Options for Microspacecraft…, AIAA
Progress in Astronatics and Aeronautics Vol. 187
4 Mueller J., in “Micropropulsion for Small
Spacecraft” (M.M. Micci and A.D. Ketsdever,
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on GOCE. SP-465, ESA 3rd International
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December 2000
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