47th AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit
31 July - 03 August 2011, San Diego, California
AIAA 2011-6129
Spacecraft System Level Design with Regards to
Incorporation of a New Green Propulsion System
Nils Pokrupa*
OHB Sweden, Solna, 17104, Sweden
Kjell Anflo† and Oskar Svensson‡
ECAPS, Solna, 17104, Sweden
This paper presents the lessons learned from the design, development and in-space
demonstration of the novel High Performance Green Propulsion (HPGP) system as
implemented on the Prisma spacecraft platform. The opportunity to fly the HPGP system
served as means to flight demonstrate the new propulsion technology, but also served as a
demonstration of how to incorporate system level aspects to the spacecraft level design.
Implementation of the HPGP propulsion system impacts five main system level interfaces
namely, thermal, power, shock, vibration and plume effects. This paper presents how these
requirements were met by spacecraft design, and quantitatively discusses the interfaces that
are to be incorporated in to the spacecraft platform based on design, ground test data and
flight test data.
Nomenclature
α
A
Cg
ɛ
F
Isp

Q
r
T

ΔV
=
=
=
=
=
=
Absorptivity
Area
Center of gravity
Emissivity
View factor
Specific impulse
=
=
=
=
=
Heat flux
Axial distance from nozzle exit plane
Temperature
Cone angle to the thruster axis
Delta velocity
*
Spacecraft Systems Engineer, OHB Sweden, [email protected], Solna Strandväg 86.
Chief Engineer, ECAPS, [email protected], Solna Strandväg 86, AIAA Senior Member.
‡
Engineer, ECAPS, [email protected], Solna Strandväg 86, AIAA Regular Member.
1
American Institute of Aeronautics and Astronautics
†
Copyright © 2011 by OHB Sweden. Published by the American Institute of Aeronautics and Astronautics, Inc., with permission.
I. Introduction
H
igh Performance Green Propulsion (HPGP) technology is based on an Ammonium DiNitramide (ADN)
monopropellant.1,2 It has now successfully been flight demonstrated and operated for over two hours of total
burn time on the Prisma mission.2 HPGP stands to be a logical replacement for hydrazine based systems since the
performance is higher yet the system is built on the same off the shelf hardware as conventional hydrazine systems.
The major advantage is in the propellant handling and loading of an HPGP propulsion system, which is reduced risk
and time1. In comparison to hydrazine, the HPGP system requires higher pre-heating power and operates at
significantly higher combustion temperature, guidelines to handle these differences, with respect to the spacecraft
level design, are given in this paper.
Prisma is a pair of Swedish technology test bench spacecraft, focused on the areas of new formation flying and
propulsion technologies.3 Launched in June of 2010, they have been in full separated operation since August of
2010 demonstrating autonomous formation flying, rendezvous and proximity operations using a suite of formation
sensors.4 The larger of the two spacecraft, Mango, carries with it three propulsion systems. Of these, hydrazine is
used as the baseline system, where as the HPGP system in addition to provide the ΔV required for the formation
flying manoeuvres is flying as flight demonstrator.
2
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II. Prisma Mission Overview4
The Prisma smallsat in-orbit test-bed was launched on the 15th of June, 2010 to demonstrate strategies and
technologies for formation flying and rendezvous and to serve as a test bench for a number of new technologies,
including the HPGP system. The mission consists of two spacecraft: Mango and Tango. Mango is 3-axis stabilized
and is equipped with a propulsion system providing full 3D orbit control capability. Tango has a simplified solar
magnetic control system and does not have any orbit control capability. The two spacecraft were launched clamped
together into a 720/780 km altitude sun synchronous dawn-dusk orbit, and later separated in August of 2010. Since
then, the two spacecraft have been performing a steady march through a tight mission and experiment timeline.
OHB Sweden (formally SSC Space Systems Division) is the prime contractor for the project which is funded by
the Swedish National Space Board (SNSB) with additional support from the German Aerospace Center (DLR), the
French National Space Center (CNES) and the Technical University of Denmark (DTU).
Figure 1. Prisma space segment summary
The figure above is a summary of the space segment hardware that constitutes the Prisma mission. The above
space segment hardware is thus intended to support the defined mission to demonstrate formation flight and
rendezvous, whilst also providing “first flight” opportunities for a number of new sensor and actuator technologies.
Thus the demonstrations can be divided in to Guidance, Navigation and Control (GNC) Experiments and Hardware
Tests.
III. Prisma Implementation of HPGP
A. Performance
For continues firing with near Steady-State
conditions the improvement in HPGP Isp over
hydrazine is 6% at BOL and 12% towards EOL. For
Single Pulses the improvement in HPGP Isp over
hydrazine is 10% at BOL increasing to 20% towards
EOL. For Pulse Mode at very low duty and low
propellant feed pressure, the HPGP performance is
comparable to hydrazine performance. At some
pulse modes the HPGP performance has up to 12%
higher Isp with hydrazine. Detailed information with
respect to in-space demonstration of the HPGP
system is given in Ref. 2.
Figure 2. PZ 1 Newton HPGP thruster assembly
(payload camera baffle in upper right background)
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B. Mechanical Interfaces
The PRISMA HPGP system consists of one diaphragm-type propellant tank with a capacity of 5.5kg (i.e. 4.5 L)
of LMP-103S propellant, two service valves, one pressure transducer, one system filter, one isolation latch valve and
two 1 N thrusters. The propellant and the pressurant gas are stored in the tank and are separated by means of a
diaphragm. The pressurant (Helium) acts on the flexible diaphragm and pushes the propellant via the system filter to
the thruster propellant Flow Control Valve (FCV)5.
Pressurant Service
Valve
PZ HPGP assembly
GHe
LMP-103S
Propellant Service
Valve
Orifice
Pressure
Transducer
Filter
Latch Valve
TS
TS
MZ HPGP assembly
Thrusters
Figure 4. Thruster positions on Prisma
The hardware associated with an HPGP
system is exactly that of a typical Hydrazine
system, with two exceptions, as outlined in Table
1 below.
The majority of the hardware was mounted
with standard interfaces and procedures as a
typical hydrazine based system. This includes the
supporting the tubing on isolated standoffs,
wrapping tubing with low emissivity tape, five
layer Multi Layer Insulation (MLI) around the
tank and standard vibration precautions. Due to
the high operational temperatures of the thruster
chamber during firing, some additional care has
been taken on the interface and isolation of the
thruster from the spacecraft. The section III-C
below discusses these requirements in detail, but
the resulting mechanical isolation and bracket
layout is shown in the images below.
Figure 3. Hydraulic schematic
HPGP system on Prisma5
for
the
Figure 5. Prisma Mango, flight configuration, during
environmental solar simulation testing
Table 1. System hardware list
Tubing
Service valves
Pressure transducer
Filter
Latch valve
Orifice
Tank & diaphragm
Flow control valve
Thruster
COTS*






Re- Qualification
New Design



* hydrazine based Commercial Off The Shelf hardware
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M5X25_MC6S (x3)
Thruster mounting plate
M5 Thermal Bushing (x3)
Washer (x3)
M05 (x3)
Thermal standoff
FCV
Bracket bottom
Bracket lid
Figure 6. External mounting description
Figure 7. Internal mounting description
The sun facing elements of the mounting hardware were covered in high temperature 10 layer MLI (qualified to
peak temperatures of +350oC and +220oC continuous operation). Other surrounding surfaces, such as spacecraft
panels, >10cm from the thruster chamber used nominal 10 layer MLI (qualified to +150 oC). Each bracket was
provided with approximately 10x10cm of radiator surface area to cool the bracket hardware and FCV (consisting of
silver Fluorinated Ethylene Propylene (FEP) tape, ɛ=0.76, α=0.05). The thermal requirements are discussed in detail
in Section III-C below.
The resulting vibration environment subjected on the two HPGP thrusters during Prisma spacecraft level flight
acceptance testing are shown below in the coloured data (PZ and MZ thrusters, in x, y, and z axes). Also shown (in
red) are the unit level acceptance and qualification specifications for the thruster assembly against which the thrust
assemblies have been verified.
Figure 8. Random vibration at two thrusters I/F during acceptance (lateral on left, axial on right)
The acceptance level spectrum corresponds to a 16.14 gRMS energy, which is derived from the ESA “ECSS-E10-30A Space Engineering” standard based on a unit mass of 0.378kg. The resulting responses seen by the thruster
during acceptance testing (shown above) corresponds to 18.49 gRMS lateral and 4.69 gRMS axial loads.
Figure 9. Shock test response data (Y lateral, Z lateral, X axial)
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C. Thermal Interfaces
1. Thrust chamber temperature
It is clear that one of the major factors
for spacecraft level integration of the HPGP
system is the high operating temperature of
the catbed and combustion chambers.
During firing, the combustion chamber
reaches quickly 1480oC (see the profile in
Figure 10 below). Prior to firing, the catbed
temperature is regulated by software
between 340oC to 360oC, which is the
nominal start temperature. When the firing
command is given the catbed heater is
switched off during the duration of the pulse.
Figure 10. Thrust chamber skin temperature at steadyIn-space data shows that for short pulse
state firing
durations (≤100ms) and a duty of
approximately 1%, the temperature of the
Table 2. Heat dissipation from thruster
catbed initially drops ~20oC. This has not
Heat Dissipation per 1 N HPGP Thruster
lead to any malfunction, however for future
missions it is recommended that the thermal
Pre-Heating
Steady-State Firing @ BOL
control of the catbed heaters are independent
from the firing command and hence actively
Conduction*
Radiation
Conduction*
Radiation
maintining minimum required temperatures.
2.3 W
7W
2.5 W
207 W
Due to the high temperatures particular
*
Conduction
to
spacecraft
via
thruster
mounting
bracket
thermal isolation measures are taken,
including the mechanical thermal standoff
and bushings described in section III-B. The Prisma mission has taken the opportunity of various duty cycles, pulse
lengths and environmental loads to qualify the operation of the thruster and FCV successfully in all cases 2.
The FCV must be maintained between 10oC and 60oC at all times during operation, although the operational
temperature of the system is between 10oC and 50oC.
The propellant ignition, decomposition and combustion occurs in the catalytist bed (ie: catbed, which is the
thermocatalytic reactor).
2. Time required for pre-heating
The time required for 9.25W of catbed
preheating (to a minimum of 340oC) has
been shown in flight to be roughly 600s to
720s,
irregardless
of
the
initial
environmental conditions.
The data in
Figure 11 from 18-08-2011, for instance,
shows initial temperatures of 74oC and 32oC
resulting in diffent sunlit conditions. This
matches with ground test performance.
The preheating time is set conservatively
to 30 minutes before enabeling the HPGP for
firing. However the thruster reaches the
nominal start temperature after 12 minutes of
preheating. The average temperature of the
HPGP propellant tank was maintained at the
nominal temperature of 20 to 23oC.
Figure 11.
Time required for pre-warming of catbed heaters
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Of note, the data shown in Figure 11 is for regulation between 360oC and 380oC, this was later lowered to 340oC
to 360oC during the nominal mission. The duty cycle of the catbed heaters during regulation has been quantified to
between 67% and 80%.
(A)
(B)
(C)2
Figure 12. FLIR image during preheating (A) and continuous firing (B), with visual image during firing (C)
3. Soak back effect
When firing a thruster there is a thermal
soak back after each pulse (with the exception
for very short pulses with low duty factors) of
residual heat that flows back from the thrust
chamber assembly towards the thruster
propellant flow control valve. The impact of the
soak back on system level has been measured
for the HPGP system and quantified in flight.
Figure 13 shows flight data from six orbits
during 23-08-2011 for typical soak back curves
on the FCV during a set of two PZ and two MZ
firings, all within expected ranges. When
detectable, the change in temperature is <5oC.
The details of Figure 13 are the following:
* - Each of these increases correspond to the
activation of the tubing heater lines, so not
related to catbed heating or firing.
A – 25% duty pulse train on MZ
B – 25% duty pulse train on PZ
C – 20s single pulse on PZ
D – 20s single pulse on MZ
E – The cycling in PZ FCV temperature is
due to the change in solar angle (bottom figure)
as a result of rotations exposing the FCV
mounting panel to sun. This is not the result of
soak back effect.
Figure 13. Typical soak back curves on the FCV and
spacecraft relative sun vector
The “MZ FCV” shows the most typical soak back effects, all <5oC in temperature change (points A and D). In
the case of the PZ FCV the soak back effects are overridden by sun angle effect on the spacecraft interface as the
valve mainly follows the spacecraft body temperature (points B and C, see III-C-5 on thermal strapping).
Fuel line temperatures also change, as expected, due to the temperature of the fuel released from within the tank
(roughly 23oC), thus a slight temperature rise is seen on the colder MZ line and slight temperature decrease on the
warmer PZ line.
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4. Surrounding MLI
Hot firing ground tests have been
performed with nominal 10 layer MLI
(+150oC max) in the chamber during long
firings (30 to 300 seconds) on 12 May 2009.
The photo below shows the MLI and
thermocouple placement after these tests. It
is clear that even the nominal 150oC MLI
survived, although one can see heat ring
markings around the thruster up to roughly
30mm. The thermocouple was placed in this
region and measured a max temperature of
+126oC during the longest firings. As a
result of these tests, the high temperature
(+350oC) MLI was applied closely around
the thruster in flight.
The star camera MLI is rated to +150oC,
and considering the star camera MLI is
nearly 250mm away from the thruster, it can
already be safely assumed that the star
camera MLI will not encounter overheating
issues. Thus the black surfaces in the image
below were covered only with nominal
(150oC) MLI.
Figure 15.
~30mm
Figure 14. Test of MLI survivability when tighly wrapped
around thruster
~250m
Star camera baffles in m
proximity of the thruster
To be complete, an estimation using Eq. (1) and Eq. (2) of the radiative exchange between the thruster-MLIspace can be made in order to quantify the results, although a number of assumptions have to be made:
 eff 

 i j
1  Fij F ji 1   i 1   j 

Qij   eff Fij Ai Ti 4  T j4

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(1)
(2)
Table 3. Thruster to MLI radiative heat exchange parameters
Value used
Source
Tthruster
TMLI
Tspace
Athruster
AMLI
εthruster
1480 oC
-100 to +150 oC
3K
1.25x10-3 m2
0.1x0.2 m2
0.32 to 0.50
HPGP data (Figure 10)
Max operating range, typical MLI data sheet
Arno Penzias & Robert Wilson
HPGP data
One exposed side of baffle
HPGP data
εMLI
Ftm
0.85
0.06
Fmt
0.02
Fms
0.75
εeff
0.425
MLI data sheet
Thruster-MLI view factor, conservative estimate from
Thermica radiation model
MLI-thruster view factor, conservative estimate from Thermica
radiation model
MLI-space view factor, conservative estimate from Thermica
radiation model
Effective emissivity thruster-MLI, Eq. (1)
Star camera baffle MLI radiative exchange
25
20
Radiative exchange (W)
These values can then be plotted all
together using the equations above and the
ranges of 0.32 to 0.50 for thruster “ε”
(emissivity) and -100 to +150 oC MLI
temperatures, to get the following plot in
Figure 16.
From this rather conservative approach
and assumptions, it is suggested that the
resulting MLI balance temperature would
be between +80 to +120 oC (using more
realistic view factor values, these numbers
fall to +35 to +70oC). This is within the
allowed <150oC range for typical exterior
MLI.
15
10
5
Rad to space
Rad from thruster (e0.50)
Rad from thruster (e0.32)
0
-150
-100
-50
0
50
100
150
200
MLI temperature (C)
Figure 16.
baffle MLI
Thruster radiative exchange to star camera
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5. Thermal strapping of the FCV
For the initial mated checkout operations (with Tango
still attached to Mango, blocking an interface cooling
radiator) it was necessary to include a cooling strap on the
PZ FCV. The opportunity was taken to mount the two
HPGP thruster in two different configurations, with and
without thermal strapping, allowing for investigation of
thermal behaviour.
With potentially 3 to 4 W entering the FCV and
needing to be dissipated, it was clear that a 11 K/W
estimated coupling from the FCV to the spacecraft (9 K/W
to the mounting plate plus 2 K/W to the spacecraft panel),
signficatly large temperature increases could be seen on
the FCV (~3W*11K/W = 33K temperature rise). A more
acceptable temparutre rise is in the order of 10oC, which
would require a FCV to spacecraft coupling in the order of
<4 K/W. One option is to thermally strap the mouting
plate to spacecraft interface, or better directly strap the
FCV to the spacecraft structure (see Figure 17).
Thus the PZ thruster was outfitted with a copper
thermal strap to the panel, effectively coupling the FCV
temperature to the spacecraft panel temperature of 35 to
40oC (see Figure 18). Whereas the MZ thruster was not
strapped and only relied on conductance through the
screws and mounting bracket and radiation.
The difference in performance between the PZ and MZ
FCVs can be seen during the activation of the catbed
heating, shown in Figure 19. Prior to catbed preheating,
both valves are sufficiently cool that their heater control is
active (regulating between 18 and 21oC). With the
activation of the 360 to 380 oC catbed regulation (later
reduced to 340 to 360 oC), the valves warm up, but it is
clear that the thermally strapped PZ FCV settles and only
increases 10 oC in temperature. Whereas the MZ noncooled valve increases nearly 30 oC in temperature, which
agrees with original estimations mentioned earlier.
The trapping of heat in the valve also became evident
during other more demanding orientations and pulse trains
in which it was clear that the MZ valve temperature was
steadily increasing. The nominal sun vector plot confirms
that this was not an effect of environmental loads. This
was not the case with the thermally strapped PZ FCV, thus
it is clear that a sufficient heat path of <4 K/W to a
spacecraft sink of <40oC must be provided to the FVC
hardware.
The resulting mounting interface temperature
requirement for the FCV+thruster assembly is then 0 to
35oC, to allow for +10oC dissipation, +5oC soakback and
margins.
Depending on spacecraft mounting
configuration, this may be increased to 45 oC.
Figure 17.
FCV thermal strapping options
Figure 18.
PZ installed thermal strapping
Figure 19. PZ & MZ FCV temperature response
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D. Power and Control Interfaces
The interface to the Prisma Data Handling System (DHS, aka: on board computer) is via interface cards such that
the interface to the spacecraft is simply via a CAN Bus and 28V power supply. This can also configured to interface
to an RS-422 bus, if required on other missions. A schematic of the fully redundant (A+B) Prisma configuration is
shown below.
Figure 20.
Prisma control interface schematic
Figure 21. Example of upgraded control
electronics configuration (Proba-3 mission)
A slightly alternate interface configuration upgrade has been proposed for the Proba-3 mission, as shown below.
Proba-3 is the next ESA formation flying mission, in which HPGP has been baselined. This includes the following
interfaces and functions:
• 8 Supply outputs for Thruster valve control.
• 2 Bi-polar outputs for Latch Valve control.
• 2 Inputs for Latch Valve status monitoring, to avoid failure propagation, sensing the latch valve status
i.e. closure, uses high impedance input sensing
• 2 Dual diode short circuit protected 20V Pressure Transducers supply
• 2 Inputs for Pressure Transducers acquisition
• 8 Inputs for Thermocouples of type N or K.
• 8 Outputs for HPGP catalytic bed heater.
Thermal control of the tanks, tubing and FCVs is not included in the control electronics as these are of the
spacecraft platform responsibility. These are typically surface mounted foil heaters, controlled by software setpoints
and in some cases protected by an overheat thermostat. The interface cards do, though, read in the catbed
thermocouples and drive the catbed heaters.
Catbed pre-heating is individually regulated by software between 340 to 360 oC (i.e. Thermocouple
Temperature) for each thruster. During firing the catbed heating is disabeled when the thermocouple temperature
increase over 360 oC and is ”OFF” until the temperature decrease under 340 oC. The FCVs were fitted with heaters
regulated by software between 18 and 21 oC, although these were only required on a 50% duty cycle prior to the
activation of the catbed heaters.
As a result of these various heating and operating modes, the following FCV and thruster power consumption
table has been compiled from Prisma flight data. Not included in these figures is the additional 0.1W required to
operate the drive electronics interface card.
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Table 4. In flight power consumption summary
Power Consumption per 1 N HPGP Thruster @ 28 VDC [Watts]
FCV Heater
FCV
CatBed Heater
Nominal Avg. Peak Max Min Peak Max Min Peak Max Min
1.5 W
3.0
3.0
0
0
0
0
0
0
0
System Disabled, Pre-heating
“OFF”
7.3 W
0
0
0
0
0
0
9.25 8.0 6.75
Pre-heating or System
Engabeled Pre-heating “ON”
1.5 to 9.8 W
0
0
0
8.3
8.3
1.5
8.3
8.3
0
Firing
-
signifies functions that are not used or disabled in the mode of operation
E. Environmental Interfaces
Clearly the preheated catbed temperature varies based on the relative sun vector on the thruster body. This can be
seen in the range of heater power values in the table above, as these fluctuations are driven mainly by additional
solar flux. By collecting the flight thruster heater performance data and plotting vs the sun vector, a relationship
summary can be made of the required catbed heater power vs the solar flux and view to cold space. This has been
simplified in the figure below.
Figure 22.
Pre-heating power consumption, wrt to environmental loads
The 10W catbed heater is clearly not saturated in any of the environmental cases. The lower values of 6.7 to
8.0W is simply a result of the 340 to 360 oC temperature regulation, implying a 67 to 80 % duty cycle on the catbed
heater. Logically, the lowest duty cycle (67%) occurs when the solar flux is perpendicular to the thruster body
(largest projected area to the sun), and the highest during cases of no illumination such as eclipse. The surface
properties of the main exposed thruster body are Rhenium, and thus have properties of ɛ=0.32 to 0.50 and α=0.42.
Prior to catbed activation during commissioning, it had been seen in the flight data that the temperature of the
thruster chamber could easily change by +/- 60 to 80oC depending on sunlit conditions.
F. Launch preparation logistics
Although beyond the scope of this paper, it is worth mentioning one of the key system level advantages to the
HPGP technology which is the fuelling and decontamination process. On ground logistics is simplified by orders of
magnitude as a result of the HPGP classification as an insensitive substance (NOL 1.3). As the system was declared
as a “Non Dangerous Operation” by the Yasny Launch Base Range Safety, all other activities such as launch
preparation of the other satellite PICARD could continue without any restrictions during the HPGP fuelling process.
The quantitative details of the fueling and decontamination process for HPGP vs hydrazine (crew size, number
of days for the procedure, waste material, etc) has been presented in Ref. 2 and Ref. 6, and will not be further
elaborated in this paper.
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G. Plume effects
No plume effects have been observed during the first year of operation in space. Long term mission data has not
reveiled any contamination or degration effects on solar panel or camera performance.
1. Plume analysis
A plume impingement analysis for the
max
Prisma spacecraft was performed7 with the
r (m)

objective to estimate the maximum heat load
on the solar panel when the 1N HPGP thruster
are fired.
The model used work follows closely the
work of Boettcher and Legge who developed
the DLR computer code for plume analysis
and impingement applications. The work of
Boettcher and Legge13-15 is built on a wellestablished theory for expansion of a nozzle
flow into vacuum, and essentially the same
modeling is presented in the literature by
several authors8-16. The model structure
consists of two major parts: A first part,
hereafter referred to as plume analysis, which
models the structure of the plume as it exits the
nozzle and expands into vacuum. The second
step, hereafter referred to as impingement
analysis, is to determine the heat flux a surface
located such that it is exposed to the expanding
Figure 23. Plume axis definition
plume.
The plume analysis is a far-field approximation of a supersonic plume expanding into a vacuum. The far-field
approximation builds on the assumption that the streamlines of plume, far from the nozzle exit, are straight lines
forming a cone with the ½ opening angle max (15 deg), inside which the plume is contained. Another important
assumption is that the density along the plume’s axis of symmetry varies as 1/r 2, where r is the axial distance from
the nozzle exit plane. With these assumptions, it is suitable to model and describe the plume in a polar coordinate
system of r and θ. With these assumptions, it is suitable to model and describe the plume in a polar coordinate
system of r and θ as shown in Figure 23.
2. Impingement analysis
The impingement analysis provides an estimate of forces and heat loads, and provide input regarding
contamination, on the given surface.
To estimate the surface’s heat flux with the present model, the following input is needed:
• Surface temperature
• Surface orientation, i.e. the angle between the surface normal and the impinging velocity vector
• The surface’s energy accommodation coefficient.
The objective of the analysis presented here is to estimate the effects of plume impingement, primarily heat flux,
from the 1N HPGP thrusters onto the spacecraft. The geometry and thruster location for the Prisma spacecraft is
shown in Figure 4 Schematic view of Prisma with locations of star cameras 2 HPGP thrusters visible. The most
critical parts, from a plume impingement point, is the rear side of the solar panel, star cameras, payload cameras and
one of the payload antenna. Star camera baffles which intersects with the thruster plume in a plane that is parallel to
the plume axis but shifted approximately 400 mm from the same. In polar coordinates the most critical point of
impingment is on the line between (r,ax).
The star cameras will partially stare through the HPGP thrusters plume at a location more than 300 mm from the
thrusters exit plane. i.e. in polar coordinates: r > 300 mm.
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3. Thruster operations and characteristics
During the Prisma mission, both continues and pulse mode firings of the HPGP thrusters are performed.
Continues firings are performed up to 60 seconds during manoeuvring. As a worst case assumption, the plume
analysis reported herein, corresponds to contiunues firing only. Pulse mode operation is regarded a less severe case
in every respect. Up to three thrusters are fired at a time on Prisma, but due to the thruster locations, effects of
plume/plume interaction are not considered. The following thruster and propellant characteristics are used for the
analysis:
Thruster
Gas dynamical properties of the exhaust gas from
LMP-103S
- Combustion chamber stagnation pressure: 15 bars
(absolute)
- Nozzle exit area: 38 mm2
- Nozzle divergent half angle: 15º
- Nozzle expainsion ration: 100:1
Flow ≤ 0.5 g/s
Exhust spieces composition (mole fractions):
-
50% H2O
23% N2
16% H2
6% CO
5% CO2
Spacific heat ratio k = 1.23 (-)
Specific heat at constant pressure 2.284 (kJ/kg-K)
Molecular mass = 19.7 (kg/kg-mol)
Gas constant per unit weight R = R´ / Ӎ (J/kg-K)
Universal gas constant R´ = 8314.3 (J/kg mol-K)
 (deg)
r (m)
Figure 24. 1 N exhaust plume Density Field [kg/m3] (in polar coordinates)
In polar coordinates, the most critical point of impingment is somewhere on the line between
(r,and with units in mm and degrees.
The star cameras will partially stare through the HPGP thrusters plume at a location more than 300 mm from the
thrusters exit plane. i.e. in polar coordinates: r > 300 mm.
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 (deg)
r (m)
Figure 25. 1 N exhaust plume Temperature Field [K] (in polar coordinates)
 (deg)
r (m)
Figure 26. 1 N exhaust plume Heat Flux Field [W/m2] (in polar coordinates)
The result of this study indicates a maximum heat flux value of 62 W/m2 on the rear of the solar panels. On all
other surfaces the impinging heat flux from the HPGP thrusters is significantly less. This is considered safe, as
compared to solar flux, for instance, which is 1350 W/m2.
The star cameras will partially be staring through a thrusters plume
with the following main characteristics:
• Density: < 10-6 kg/m3
• Temperature: < 70 K
A series of hot firing tests on ground using withness plates and post
test Scanning Electron Microscopy analysis were performed to
investigate the total contamination in the plume < 10 ppm. The
contaminants are shown in Table 5.
The contaminants acts like particles and not as gas molecules and
therefor only travles along the core of the plume with almost no angular
dispersion.
Table 5. Plume contaminents
Contaminent
< 5 ppm
S
< 5 ppm
K
< 0.5 ppm
Ca
< 0.5 ppm
Na
< 0.5 ppm
Na
< 0.2 ppm
Si
< 0.2 ppm
Cu
< 0.1 ppm
P
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IV. Interface Requirement Summary
Figure 27. Prisma 1 N thruster assembly interface drawing
Figure 28. Prisma 1 N thruster assembly electrical interface notes
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Table 6. Prisma interface summary
Mechanical (see Figure 3 to Figure 7)
Thruster + FCV mass
Interface cards (redundant)
Mass of pressure transducer, latch valves, tubing,
service valves, tanks, standoffs, etc.
Random vibration at mounting plate
Shock
Thermal
Catbed pre-heat temperature range
Max thruster firing temperature
Operating fuel temperature
Interface temperature range at mounting plate
FCV operation range
Power consumption (see Table 4)
FCV heater power
Catbed heater power
Drive electronics power consumption
Interface to spacecraft
Data and command
Power
Heat flux to spacecraft during operation
0.378 kg
0.9 kg
(Standard COTS)
PSDmax = 1.82 g2/Hz (qualification)
= 0.455 g2/Hz (acceptance)
10,000 g @ >1000 Hz
340 to 360 oC
1480 oC
10 to 50 oC
0 to 35 oC
10 to 60 oC (with <4 K/W coupling to spacraft sink)
3.0W (max), 0.0W (avg. during normal operation)
9.25W (max), 7.3W (avg. during normal operation)
0.1W (avg. during normal operation)
CAN Bus or RS-422
28V DC, 12V DC for continuous thrusting mode
< 2.5 W conduction
< 1.5 W radiation
V. Conclusion
The information presented above has demonstrated, by means of analysis, design and one year of flight data, the
implementation of the HPGP system on-board a spacecraft platform. Key interfaces and design aspects such as
thermal, power, vibration, environment and plume interaction have been discussed and solutions to those interfaces
have been quantified.
In this regard, the Prisma design and mission has served as a successful in-flight demonstration of the system
integration and application of a new High Performance Green Propulsion system. The experience and the
knowledge gained during the Prisma mission are valuable for the implementation of HPGP in systems design for
future missions.
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Acknowledgments
This work has been performed under contract from the Swedish National Space Board (SNSB). The authors wish
to acknowledge the sustained support from SNSB, ESA, OHB Sweden and SSC. The authors also acknowledge the
strong support from the effort of all co-workers in this project from ECAPS, OHB Sweden, SSC, Royal Institute of
Technology, and EURENCO-Bofors. The authors also wish to acknowledge the DLR crew at the German Space
Operations Center (GSOC) in Oberpfaffenhofen during the HPGP 4 operations in May 2011 and the personnel at the
three Ground Stations at Esrange, Weilheim and Inuvik.
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