SP2012_2362672
The Design, Development, Manufacture and Test of an Electric Propulsion System for
Small Spacecraft
Mark Pollard
Surrey Satellite Technology Ltd
Guildford, United Kingdom, m.pollard@sstl.co.uk
Dimitrios Lamprou
Surrey Space Centre, University of Surrey
Guildford, United Kingdom, d.lamprou@surrey.ac.uk
successful RapidEye, Proba-2 and NigeriaSat-2
spacecraft. The propulsion system acted as a test
bed for new processes and techniques, and this
included the development of a Titanium Metal
Matrix Composite tank by TISICS Ltd, although
this was not subsequently integrated for this
mission.
Abstract
This paper discusses the design, development,
manufacture and test of the electric propulsion
system manufactured by SSTL for the UK
technology
demonstration
spacecraft,
TechDemoSat-1 (TDS-1). The spacecraft is funded
by TSB and SEEDA to serve as an orbital test bed
for the emerging technologies from the UK space
industry.
Acronyms
AIT
ALM
BN
COTS
CoG
EVT
FM
HCT
MEOP
MMC
OBC
PPU
QM
TechDemoSat-1 does not require a propulsion
system to complete the mission; however SSTL
opted to design and build the system with internal
funds to test new technologies and processes. The
SSTL design approach has proved valuable for this
task, with the spacecraft schedule lasting less than
two years between concept and launch.
The TDS-1 propulsion system is equipped with
two thrusters; one is an SSTL resistojet and the
other is a Hollow Cathode Thruster (HCT)
developed by the SSC. This complimentary
approach of a resistojet and an efficient electric
thruster is an important development allowing
small spacecraft to achieve more complicated
missions. These two thrusters are useful together,
with the resistojet providing a higher thrust for
shorter manoeuvre times and the HCT providing
low thrust and more efficient operations. Whilst
the design of the HCT is not currently optimised, it
promises to provide good propulsive efficiencies
for small power limited spacecraft and forms an
essential component in several other efficient
electric propulsion devices in development at the
SSC. The development of this unit has included the
control electronics, which are also designed to be
applicable to a variety of electric thrusters
currently in development.
SEEDA
SSC
SSTL
TDS-1
TIG
TSB
Assembly, Integration and Test
Additive Layer Manufacturing
Boron Nitride
Commercial Off The Shelf
Centre of Gravity
Environmental Test
Flight Model
Hollow Cathode Thruster
Maximum Expected Operating Pressure
Metal Matrix Composite
On-Board Computer
Power Processing Unit
Qualification Model
South East England Development
Agency
Surrey Space Centre
Surrey Satellite Technology Ltd
TechDemoSat-1
Tungsten Inert Gas
Technology Strategy Board
TechDemoSat-1
The TDS-1 spacecraft is a technology
demonstration satellite based on SSTL’s highly
capable SSTL150 platform, which formed the core
of the RapidEye constellation. TSB and SEEDA
have funded the heritage platform and mission, and
companies and institutions across the UK are
supplying payloads at their own cost in order to
obtain flight heritage and display the novel space
The feed system is based on the heritage SSTL
xenon feed system as flown on the highly
1
technologies available in the UK. The
developments on board range from radiation and
cosmic ray detectors, to a new on-board computer,
to a de-orbit device being developed to help reduce
space debris, and a payload to monitor shipping
and the maritime environment. A rendering of the
TDS-1 spacecraft can be seen in Figure 1.
•
•
•
•
Qualify a new tank weld geometry
Qualify a new system welder
Develop
an
in-house
cleanliness
verification facility
Develop the test and verification approach
required for electric propulsion systems
Design and System Architecture
The propulsion system design is largely based on
SSTL’s heritage xenon propulsion system as used
on Beijing-1, RapidEye, Proba-2 and NigeriaSat-2.
A schematic of the propulsion system can be seen
in Figure 2. It comprises an electronic bang-bang
regulation system, using feedback from the low
pressure transducers, to control series-parallel
solenoid valves in order to reduce the nominal
80barA tank pressure to a nominal 1barA plenum
pressure. This is used to feed a resistojet and the
HCT and can be adjusted to feed both thrusters as
described below.
Figure 1: TDS-1 Spacecraft
SSTL is able to utilise the platform to test new
product developments, and the propulsion system
is one such example. SSTL and the SSC have
invested in the system to prove the complimentary
capability of a high thrust resistojet and a low
thrust, higher efficiency electric thruster and test
new processes and technologies. This will help
enhance the capabilities of small propulsion
systems and enable new missions for small
satellites.
Propulsion System Requirements
The following is a list of internal requirements that
were placed upon the propulsion system:
• Adapt and subsequently prove the
compatibility of a typical SSTL feed
system with an electric thruster
• Prove the complimentary operation of a
high thrust resistojet alongside a high
efficiency electric thruster being fed by a
common feed system
• Advance the HCT to a flight status
• Advance the HCT PPU to a flight status
• Gain flight heritage on the HCT and
electronics (also utilised by other electric
thrusters in development at the SSC)
• Accurately characterise the performance of
the HCT during in-orbit operations and
correlate with ground test data
• Qualify a new tank manufacturer
Figure 2: TDS-1 Propulsion System Schematic
With two thrusters being fed, thruster isolation
valves were required. The HCT also required a
reduced flow rate, compared to the resistojet,
which was achieved using a COTS sintered metal
2
flow restrictor press-fit in a manifold. The system
had to fit in the volume of the original SSTL150
propulsion system. Due to the additional valves,
pipework and electronics required to support the
HCT, a new tank and a new system design was
required.
Tank
With the system volume defined by the available
space in the platform, the tank was made as large
as possible whilst allowing the system to fit in the
space available. This provided a tank which could
hold 1.75kg of xenon propellant. This would allow
enough propellant to prove the HCT and collect
performance data, whilst allowing a useful
propulsive capability to be available for the
spacecraft. The tank was designed with operator
safety as the most important factor, and due to the
condensed schedule it was essential the design did
not fail during testing. This has resulted in an overdesigned tank, but one that is compatible with a
tight schedule and low cost.
There are two independent propulsion controllers,
with propulsion controller 0 responsible for the
resistojet string and propulsion controller 1
responsible for the HCT string. In failure scenarios,
both controllers can be operated at the same time
meaning either controller can regulate the pressure
from tank to plenum for either thruster. The
propulsion feed system can be seen in Figure 3.
Heritage components were used where possible.
These included a fill/drain valve manufactured by
SSTL from a COTS Swagelok component, a
COTS Swagelok filter, COTS temperature sensors,
and aerospace quality pressure transducers
procured from Kulite Sensors.
SSTL decided to use the opportunity to qualify a
new tank manufacturer, and to qualify an improved
tank weld geometry. The tank was machined from
two solid billets of titanium 6Al-4V and these were
joined using an electron beam weld. A
qualification tank was produced and subjected to a
test campaign which included vacuum cycling,
Maximum expected operating pressure (MEOP)
cycling, proof pressure cycling, leak checking and
pressurisation to >4 times MEOP. The
qualification tank passed all of this testing and was
shown to be suitable for use. The tank will undergo
vibration at spacecraft level.
The opportunity also arose to work with TISICS
Ltd to develop a tank compatible with the system
but manufactured from Titanium metal matrix
composites (Ti MMC), rather than the standard
monolithic titanium tanks usually flown. The
primary aim of TDS-1 is to demonstrate new
technologies in the UK which would have a direct
benefit to future satellite applications, and this was
one such opportunity. Future Ti MMC tanks can
provide lighter tank solutions for the same
propellant mass, or can allow more propellant to be
stored at higher pressures for the same mass and
volume of tank.
Figure 3: TDS-1 Propulsion Feed System
The resistojet used is a standard SSTL product
which uses a 30W heater to provide energy to the
propellant and increase the specific impulse
relative to a cold gas system. It is mounted on a
bracket which allows both the pitch and the yaw
angles to be adjusted. This allows the resistojet to
be aligned such that the thrust vector points
through the spacecraft mission average CoG.
TISICS manufactured two qualification tanks and a
flight tank. The first qualification tank passed the
qualification tests which were the same as for the
standard machined tank (described earlier). The
second qualification tank was vibrated in addition
to the other testing, prior to the pressurisation to
>4x MEOP, and this also passed. However, it was
decided to integrate the standard machined
titanium tank into the propulsion system. This
decision was taken as the first qualification model
The solenoid valves used on the system were spare
valves from previous missions and were
manufactured by Marotta UK, now Ampac-ISP.
These valves have been flight qualified on the
Proba-2 propulsion system.
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The single spiral heater employed for the
laboratory HCT has been replaced by two bifilar
spiral heaters (primary and redundant). A 0.3mm
diameter tantalum wire is used for increased heater
robustness at the expense of increased current
requirement to reach temperature. Thicker wire
diameters of up to 0.5mm were tested but they
proved difficult to form in the required dimensions.
Hot to cold heater junctions are formed by swaging
the wire in 0.5mm internal diameter, 0.8mm outer
diameter tantalum tubes. The resulting conductor
has a 0.8mm outer diameter, which reduces heat
dissipation on the section of heater wire outside the
ceramic heater body and increases heater life and
mechanical robustness. Heater leads for power
return are terminated on the cathode tube by
resistive welding. Heater live wires are connected
to insulated interfaces inside the HCT by TIG
welding to bolts.
burst at a lower pressure than expected (but still >4
x MEOP) and due to the tight schedule the cause
could not be fully determined by the time system
manufacture was occurring. The project has
allowed TISICS to advance their designs and gain
some important experience and understanding, and
has brought the prospect of TI MMC tanks closer.
More information regarding the Ti MMC tank can
be found in [2].
Hollow Cathode Thruster
The HCT is being developed in a joint SSC/SSTL
venture. The HCT operation is based on field
enhanced thermionic emission. A low work
function insert is heated by an external heater to
~1100ºC. Xenon propellant supplied to the HCT is
ionised by electrons emitted from the cathode
and plasma is formed internally in the HCT
between the insert (ground or ‘cathode’) and an
anode electrode (keeper). The anode electrode is
kept at positive voltage by the PPU and current is
drawn from the insert via the plasma formed. The
discharge to the keeper electrode is self-sustained
once initiated, i.e. no heater power is needed.
The heater body is manufactured from boron
nitride (BN) AX05 as in the laboratory HCT. The
keeper is made from molybdenum, the cathode
tube is tantalum, and the ceramic insulations on the
cooler parts of the HCT are machined from BN
Grade M26 for increased mechanical performance.
A radiation shield made out of 0.025mm tantalum
foil encapsulates the heater reducing the heater
power requirement to initiate the discharge. The
radiation shield is resistive welded on the cathode
tube-heater assembly.
The HCT can provide increased specific impulse
capability for small satellites with the trade offs
being a higher power consumption and a lower
thrust than a resistojet. The HCT requires 50-60W
during heating and 40-60W during discharge at
nominal conditions. Higher power levels will be
ground-tested. Tests performed on the ground
indicate that the HCT will provide 85s of specific
impulse and 1.5mN of thrust [3]. To ensure the
thruster is accurately characterised in orbit despite
of the low thrust produced, the HCT is off pointed
from the spacecraft centre of gravity. The
performance will then be calculated based on
disturbance torques observed off of the reaction
wheels. This will provide a more accurate
characterisation of the thruster performance
compared to measuring performance from the
orbital change exerted on the spacecraft.
Figure 4: FM HCT (right) and QM1 HCT (left)
The HCT on TDS-1 is an iterative development
based on the SSC laboratory model that has
undergone testing and performance evaluation [3]
[4]. The thruster operates in the enclosed keeper
configuration with a keeper/cathode separation of 1
mm and utilises a porous tungsten insert
impregnated with earth metal oxides, often referred
to as a type S insert. The FM design ensures
survivability during launch and maximises thruster
operating lifetime.
The QM1 and FM HCT’s can be seen in Figure 4.
They are shown in their assembled state with the
feed pipes and fittings attached. The flight harness
is shown but has not been attached to the electrical
terminals. Thermocouples are attached at various
positions for ground testing to characterise the heat
loss.
The conductive path of the HCT to the spacecraft
is minimised by use of Ti 6Al-4V brackets. The
4
bracket interfacing directly to the HCT is provided
by EADS Innovation Works. It has been
manufactured using Additive Layer Manufacturing
(ALM) giving flight heritage to a component made
by this manufacturing process. An intermediate
wedge between the ALM HCT bracket and the
spacecraft is also made from Ti 6Al-4V in order to
ensure low thermal conduction to the spacecraft
during HCT operation and high stiffness during
launch.
various back up modes to be employed only in the
case of component failure.
Figure 5 shows a graph of the bang-bang pressure
regulation with the system reducing the tank
pressure of 80 BarA down to an average plenum
pressure of 1.5 BarA. The plenum pressure limits
for this test were 1.0 BarA lower limit and 2.0
BarA upper limit.
Manufacture
As per SSTL’s standard low cost, heritage
approach, the system was manufactured using hand
TIG welded joints and 37 degree AN flare fittings.
The tubing used is commercial grade stainless steel
tubing. As the TDS-1 propulsion system is a test
bed, a new supplier was used and qualified for the
welding.
As part of investment at SSTL, it was decided to
bring the precision cleaning of propulsion
hardware in house to lower costs and provide more
flexibility and control during the manufacture of
our systems. The TDS-1 propulsion system was
used as a test bed to establish the procedures
required and successfully prove the set-up.
Figure 5: Bang-bang pressure regulation
With the feed system operation confirmed, the feed
system and propulsion controller were temporarily
integrated with the spacecraft to allow the
operation using the OBC to be verified at an early
stage.
Propulsion System Test Approach
The testing was split into several elements:
• Component level testing
• Propellant feed system and resistojet
testing (including propulsion controller)
• HCT and HCT PPU testing
• HCT and feed system integrated testing
• Propulsion system check-out at spacecraft
level
HCT and HCT PPU Testing
Three HCTs were manufactured to provide a flight
model (FM) and two qualification models (QMs).
One of the QMs acts as a flight spare. The QMs
were used to characterise the thruster performance
given a range of input parameters, prove the design
through vibration and shock tests, and prove the
performance did not change pre and post vibration
to qualification levels. The FM HCT was subjected
to an acceptance test including performance
measurements pre and post acceptance vibration.
This allowed the performance envelope of the FM
to be correlated to the QM performances, allowing
less firing time to be accumulated on the FM at
ground testing.
Due to the low vacuum requirement for HCT
operation, the HCT is not fired at spacecraft level
EVT. Risk mitigation activities during the
spacecraft test campaign include propulsion system
communication with the propulsion module and
HCT PPU processors and HCT PPU health-check
using dummy loads.
Propellant Feed System
The HCT PPU was debugged and verified to work
effectively as a power supply for the HCT starting
sequence and steady state operation using dummy
loads. For the topology of the HCT PPU, see [3].
The following verification tests were performed on
The propulsion feed system was tested on its own,
with full functionality testing to verify pressure
regulation for both the resistojet and the HCT. This
included characterising the feed system at a range
of tank pressures, and testing the system using
5
the HCT PPU each time a component modification
was performed and before the HCTs were fired:
• High Voltage supply (150V) was verified
to operate at 150V constant voltage mode
and 0.5A constant current mode for the
start-up of the HCT
• HCT start-up spike on the High Voltage
line was tested by shorting live to ground
• Heater Supply was verified to operate at
18V constant voltage mode and up to 6A
constant current mode
• Low voltage supply was verified to operate
at 30V constant voltage and up to 6A
• constant current
• Low voltage supply and heater supply were
verified to work in series
• All PPU switches were checked
• Telemetry and telecommands were verified
to be received and acknowledged
• Data logging achieved on PC with a CAN
card
These series of tests mitigated the risks to the
integrated testing of the HCT and HCT PPU with
the rest of the propulsion system of TDS-1.
HCT – Propulsion Module Integrated Testing
Integrated testing of the HCT with the propulsion
module is performed at the SSC vacuum chamber
facilities and simulates as accurately as possible
the operation of the propulsion system while in
flight. An important modification in the firing
sequence between ground and orbit is the time that
the propellant starts to flow. During flight
operation of the HCT, the propellant flow starts 1020s before the discharge starts, during the heating
of the HCT. In ground testing the propellant flows
before the heating of the HCT starts to reduce the
risk of poisoning the insert. Several propellant
flows between 10 and 25sccm are to be tested to
ensure that the HCT can start under any nominal
conditions. Lower flow rates will be tested on the
QMs using a laboratory set-up.
The ability of both the prime and redundant heaters
to start the HCT was tested on both QMs. The start
time averaged at 300s with a +/-5% variation. The
HCTs were then fired using laboratory power
supplies to fully characterise the performance.
The objectives of the integrated tests are as
follows:
• Verify tank purging and filling procedures
• Correlate the flight system flow with the
laboratory xenon flow controller
• Confirm
maximum
and
minimum
regulation pressures to be used when
operating the HCT for nominal firings
• Perform firings at different flow rates to
the HCT
• Confirm the HCT characterisation with
ground equipment is equivalent to the
flight representative operation
• Characterise conductive heat loss through
HCT secondary structure
• Confirm the flight firing sequence and
scripts to be used
The HCT PPU was used to fire the HCTs in the
SSC vacuum chambers using a laboratory mass
flow controller. In Figure 6 the heating-start up and
thrusting sequences are shown. In the case depicted
the QM2 is fired with a heat up time of 330s. The
total time is 2000s (34 minutes) and various
throttle levels are commanded to the PPU. The
response of the HCT and PPU are nominal and are
annotated.
The test results of the mass flow correlation of the
propulsion feed system with the laboratory xenon
flow controller are shown in Figure 7. Due to the
much lower flow rates required by the HCT
compared to the resistojet, the regulation band to
be utilised is smaller, with a difference of just 0.10
bar between the upper and lower plenum limits.
This provides a stable flow rate for the HCT to
operate effectively. The resisojet requires a higher
flow rate so wider regulation bands are used. This
of course helps to reduce the number of cycles on
the valves. The average plenum pressure can be
modified to change the flow rate and thus the thrust
obtained, and the regulation band can be reduced
Figure 6: Telemetry from the HCT PPU during a full
firing sequence with QM2
6
or widened. As the graph shows, the system is
highly adaptable and configurable in flight to allow
a large range of flow rates to be utilised as
appropriate.
Figure 8: HCT profile and feed pressure profile during
an integrated test with the TDS-1 feed system
Figure 7: Selected plenum pressure regulation profiles
for the HCT and resistojet
For the integrated testing, the propulsion controller
and HCT PPU are to be controlled through a CAN
card and using scripts to maximise autonomy,
capture the flight system telemetry and prove the
in-orbit firing operation. To ensure the safety of the
HCT, a dry run through was conducted using the
feed system and dummy loads connected instead of
the HCT. The run through was conducted
successfully, allowing the HCT to be connected.
Several firings were then conducted using the
TDS-1 flight feed system coupled with the HCT
and HCT PPU. All firings were successful, and
were performed at three different average flow
rates to characterise the thrusters, with this data
then being correlated to data obtained during
earlier firings in the laboratory set up. A graph of
the first firing can be seen in Figure 8, with Figure
9 shows a close up of the firing. The plenum
pressure was cycled between 2.45 BarA and 2.55
BarA, providing a flow rate of ~20 sccm of xenon.
The propellant flow occurred throughout the
heating and cooling cycle for this ground test to
ensure the cathode was not poisoned. It took 51/2
minutes to heat the cathode until discharge
occurred, and this was followed by a 41/2 minute
firing. The system operated as expected in all
firings. The integrated testing was proved to be
successful and the system was shown to operate as
designed.
Figure 9: Close up of the HCT profile during the initial
firing shown in Figure 8
AIT/EVT implications
The cathode insert in the HCT is sensitive to
poisoning at elevated temperatures, unless vacuum
is of 1x10-5 mbar or less. This means the HCT
cannot be tested during AIT or even EVT when
there is a risk that the local pressure around the
insert may remain this high. Sub-system level
testing is critical to prove the system operation.
The propulsion feed system itself can still be tested
during AIT and EVT, including short resistojet
firings during the cold and hot cycles at EVT, and
by exhausting a small amount of propellant
through a cold HCT. To maximise the components
that can be tested at AIT/EVT stages, the HCT
PPU outputs will be sourced to dummy loads
simulating the various HCT firing phases. This
only leaves the HCT thruster itself which is not
7
tested during the AIT and EVT phases. This is a
standard approach for electric propulsion systems.
through AIT and EVT to test as much of the
system as possible.
Future
TDS-1 has provided a test mission to develop new
propulsion technologies and processes which will
enhance the propulsion systems built by SSTL.
The system can be advanced for future missions.
SSTL has designs for a high performance resistojet
which can achieve ~60 seconds of specific
impulse. These would allow the system to be
improved providing an even more capable electric
propulsion system for small power limited
spacecraft.
Acknowledgements
The authors would like to acknowledge the
following for their help with this project: Andrew
Devereaux, Mick Drube, David Gibbon, Alan Hill,
Dr. Sabrina Pottinger, Dr. Aaron Knoll, Theodoros
Theodorou, Lourens Visagie, Dr. Sarah Nash.
The HCT can be further optimised to improve
performance, especially in the terms of specific
impulse. Lower mass flow rates and power input to
the HCT are planned for testing. Life testing of the
HCT must also be performed, in terms of discharge
time and heater cycles. Ground testing can identify
areas of improvement in the start-up and operating
procedures to maximise the lifetime of the HCT in
orbit.
References
[1] The Design, Build, Test and In-orbit
Performance of Low cost Xenon Warm Gas
Propulsion Systems, M. Pollard, Space Propulsion
2010, San Sebastian, Spain
[2] Fibre reinforced titanium matrix composites
for xenon propellant tanks, S. Kyle-Henney, M.
Pollard, Space Propulsion 2012, Bordeaux, France
The HCT itself is an integral part of further electric
thruster developments currently being undertaken
at the SSC and the TDS-1 mission is allowing
flight heritage and data to be gained which will derisk future developments. The electronics design
will also be fed into these developments, allowing
small capable power electronics to be proven for
the family of electric thrusters at the SSC.
[3] Hollow cathode thruster design and
development for small satellites. D.Lamprou et
al. International Electric Propulsion Conference,
Wiesbaden, Germany.
[4] Impact of plasma noise on a direct thrust
measurement system. Review of Scientific
Instruments,
83,
033504
(2012);
doi:
10.1063/1.3692740
Conclusion
The TDS-1 spacecraft has provided an important
flight opportunity which SSTL and the SSC have
taken advantage of. The standard SSTL xenon feed
system has been adapted to provide propellant to
both a resistojet and a new electric thruster. The
result is a highly complementary system using a
high thrust resistojet coupled with a more efficient
electric thruster which allows small power limited
spacecraft to have more propulsion capability and
flexibility. The system can be further improved for
small spacecraft with the addition of a higher
performing resistojet and further optimisation of
the HCT. The system is also applicable to larger
spacecraft where there is more power available,
potentially allowing a resistojet to be coupled with
a higher power, higher performing electric thruster.
[5] Additive layer manufacturing for integrated
propulsion systems, S. Nash, C. Turner, P. Shaw,
V. Lappas, Space Propulsion 2012, Bordeaux,
France
[6] Design and development and In-Flight
Performance of a Low Power Resistojet
Thruster, D. Gibbon, A. Baker, D. Nicolini, D.
Robertson, C. Dye, 39th AIAA Joint Propulsion
Conference & Exhibit, Huntsville, Alabama
[7] A Xenon Resistojet Propulsion System for
Microsatellites, I. Coxhill, D. Gibbon, 41st AIAA
Joint Propulsion Conference & Exhibit, Tucson,
Arizona
The project has provided a mission on which the
limitations of having an electric thruster can be
solved, for instance the more extensive component
and sub-system level testing that was required to
prove the system, and the dummy loads required
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