END-OF-LIFE DE-ORBITING STRATEGIES FOR SATELLITES
R. Janovsky, M. Kassebom, H. Lübberstedt, O. Romberg
OHB System AG, Bremen, Germany
H. Burkhardt, M. Sippel
Space Launcher Systems Analysis (SART), DLR, Cologne, Germany
G. Krülle
DLR Consultant, Aidlingen, Germany
B. Fritsche
HTG, Katlenburg-Lindau, Germany
Active post-mission disposal of space structures will have an increased importance in the future in order
to keep the Earth orbit in an acceptable condition for the safe operation of manned and unmanned
missions. The objective of the performed European Space Agency study on End-of-Life (EoL-) de-orbiting
Strategies was to provide an overview and assessment of propulsion-related methods to de-orbit
spacecraft in LEO, or spacecraft which pass through LEO, to assess their applicability to different
spacecraft-mission combinations and to establish a know-how basis of End-of-Life de-orbit strategies. The
assessment strategy is demonstrated on the example of reference satellites.
ACS
AOCS
EoL
E.P.
GEO
H2O2
Isp
LEO
M
S/C
S/S
Nomenclature
Attitude Control System
attitude and orbit control system
End of Life
Electric Propulsion
Geostationary Earth Orbit
Hydrogen Peroxide
Specific Impulse
Low Earth Orbit
mass
Spacecraft
Subsystem
Background
The historical practice of abandoning spacecraft and
upper stages at the end of mission life has allowed
roughly 2 million kg of debris to accumulate in orbit.
According to [1], the uncontrolled growth of the
space debris population has to be avoided in order to
enable safe operations in space for the future. Space
system operators need to take measures now and in
future to conserve a space debris environment with
tolerable risk levels, particularly in Low Earth Orbit
(LEO) altitude regions.
Recent studies and publications show an increasing
probability of collisions between intact spacecraft
and debris [1], [3]. If no countermeasures are taken,
the number of debris particles will grow with a
growth rate in the order of up to 5% per year. Due to
the very high relative velocities in the order of 10
km/s, even very small particles in the millimeter size
range can destroy spacecraft subsystems and thus
eventually lead to the loss of the complete spacecraft
[2].
In principle two categories of measures to reduce
debris growth can be distinguished:
• Debris avoidance
• Debris removal
The natural cleaning of the Earth environment by the
trajectory disturbances will no longer be sufficient in
the future to keep the Earth environment in an
acceptable status for the operation of manned and
unmanned spacecraft. Simulations in [2] have
shown, that a real reduction of the debris population
can only be achieved with very far-reaching
measures. The only effective way to limit the growth
of the orbiting debris is to systematically remove
satellites and rocket upper stages at the end of their
mission life from the near Earth space.
Objectives
The work described in this paper is restricted to
spacecraft passing through the LEO region. Only
propulsive de-orbiting methods are considered. Nonpropulsive methods like drag-increase e.g. with
inflatable structures are not investigated within this
study.
The objective of this study was to provide an
overview and assessment of propulsion–related
methods to de-orbit spacecraft in LEO, or spacecraft
which pass through LEO, to assess their applicability
to different spacecraft-mission combinations and to
establish a know-how basis on end-of-life de-orbit
strategies. A major task was the identification of the
most suitable propulsion system to perform the deorbit maneuver for the different classes of spacecraft
with mass ranges from below 10 kg to more than
2000 kg. A set of evaluation criteria was established
for this purpose.
For spacecraft expected to be completely destroyed
during atmospheric re-entry with a negligible risk for
the ground population, an uncontrolled de-orbit
maneuver is permissible. In these cases it is possible
to perform a braking maneuver, resulting in a new
spacecraft orbit with a limited lifetime, which was
selected in this study to 15-40 years. This remaining
lifetime is assessed to be sufficient to provide the
above mentioned atmospheric cleaning effect. In the
case where the atmospheric destruction process is
expected to be incomplete, a controlled re-entry with
Deutscher Luft- und Raumfahrtkongress 2002
END-OF-LIFE DE-ORBITING Strategies for Satellites
R. Janovsky et al.
DGLR-JT2002-028
■ Satellites in circular Low Earth Orbits, i.e. with
prescribed re-entry location has to be carried out.
an orbit Altitude below 2000 km
Reference Satellites
■ Satellites in elliptical orbits passing the LEO,
All investigations refer to reference satellites and
missions. No restrictions applied to satellite mass.
The spectrum covered nano- and microsatellites with
masses lower than 10 kg up to large scientific
satellites like ENVISAT with masses up to some
1000 kg. For the current investigation the following
orbits were considered:
i.e. with their perigee below 2000 km
Table 1 summarises the selected reference
spacecraft, which were used for the further work in
this study. For each reference satellite, a
comprehensive data sheet was compiled, containing
all information necessary to perform the various
tasks.
Reference Satellites
Satellite Mass Category
Nano, < 5 kg
Pathfinder, m=1.6 kg
Nano, 5 - 20 kg
Munin, m=6 kg
Micro, 20 - 100 kg
Safir 2, m=65 kg
Mini, 100 - 500 kg
Abrixas, m=470 kg
Medium, 500 - 1500 kg
IRS-1C, m=1250 kg
Large, > 1500 kg
M=2420 kg
Table 1: Reference Satellites
De-Orbit Options
The principal disposal regions for spacecraft are
indicated in Figure 1.
In general, the post-mission disposal options are:
■ Option 1: manoeuvre to an orbit for which
atmospheric drag will remove the structure
within a given time frame, e.g. 25 years,
(uncontrolled de-orbiting)
■ Option 2: direct retrieval and de-orbiting,
(controlled de-orbiting)
■ Option 3: manoeuvre to one of a set of
disposal regions in which the structures will not
interfere with future space operations.
Uncontrolled De-orbiting (Option 1)
The de-orbiting is initiated by one or more
deceleration manoeuvres, either with relatively short
pulses of a “high”-thrust propulsion system or with a
multitude of low-thrust pulse- manoeuvres. The
deceleration manoeuvre sperformed in the apogeeregion of the orbit reduce the perigee altitude, thus
increasing the aerodynamic deceleration and
therefore reducing the orbital lifetime. The satellite
will be transferred into an orbit in which, using
conservative projections for solar activity and
atmospheric drag, the lifetime will be limited to no
more than e.g. 25 years.
Figure 1: Disposal regions and storage orbit
options for post-mission disposal
This is the case, if the predicted risk of human
casualties exceeds a specified limit, typically 0.01%
per re-entry event, or the spacecraft contains
hazardous objects with large masses and/or
radioactive or poisonous materials. In these cases,
the re-entry manoeuvre has to be controlled, which
means that the point of re-entry – and thus also the
point of impact on ground will have to be
predetermined. The controlled re-entry is initiated by
one or several propulsive retro-burn manoeuvres of
sufficient ∆V, in order to lower the perigee altitude
Controlled De-orbiting (Option 2)
In the case where the atmospheric destruction
process is expected to be incomplete or the residual
risk for ground population is too high, a controlled
re-entry with prescribed re-entry location has to be
carried out.
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Deutscher Luft- und Raumfahrtkongress 2002
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R. Janovsky et al.
DGLR-JT2002-028
uses a high-thrust impulsive Hohmann-type
manoeuvre, sending the spacecraft to an elliptic orbit
with 25 years remaining lifetime. The second
strategy uses a low-thrust propulsion system with
continuous operation, resulting in a spiral-type
trajectory and leaving the spacecraft at the end of the
manoeuvre in a circular orbit with 25 years
remaining lifetime.
Delta-V [m/s]
to a level which ascertains an immediate atmospheric
capture. Such a controlled re-entry under a relatively
steep atmospheric incidence angle produces a
confined ground impact area of break-up fragments
which have survived the aerothermal heating. The
ground impact area must be selected such that a
tolerable residual risk to persons on ground can be
achieved.
The principle option 3, which might be attractive
from an energetic point of view for orbit altitudes
above 1500 km with a transfer to a disposal orbit of
about 2500 km altitude and which is common
practice for GEO satellites EoL (‘graveyard’), will
not be considered here.
700
25 years, aver. A/m, Delta-V (Hohmann) [m/s] (Safir2)
25 years, aver. A/m, Delta-V (Spiral) [m/s] (Pathfinder)
600
25 years, aver. A/m, Delta-V (Hohmann) [m/s]
(Pathfinder)
25 years, aver. A/m, Delta-V (Spiral) [m/s] (Safir-2)
Low-Thrust manoeuvre
25 years, aver. A/m, Delta-V [m/s] (Munin-Hohmann)
500
25 years, aver. A/m, Delta-V [m/s] (Abrixas)
25 years, aver. A/m, Delta-V [m/s] (IRS-1C)
400
25 years, aver. A/m, Delta-V [m/s] (Rosat)
300
Impulsive manoeuvre
200
Comparison of impulsive and low-thrust manoeuvre
0
600
The main design driver for a de-orbit system is the
magnitude of the braking manoeuvre to be
performed at the end of mission life of the
spacecraft. In cases, where an uncontrolled de-orbit
is permissible, the spacecraft will be sent to a new
orbit with a pre-defined remaining lifetime, after
which the spacecraft enters Earth’s atmosphere for
an uncontrolled destructive re-entry.
Remaining Orbit Time [years]
Initial circular orbit 900 km
Initial circular orbit 600 km
40
25
15
10
Abrixas
Initial circular orbit 500 km
Initial circular orbit 400 km
0,1
Initial circular orbit 300 km
0,01
0
lifetime.graf
0,002
0,004
0,006
0,008
0,01
0,012
1000
1100
1200
1300
1400
1500
1600
1700
1800
1900
2000
Initial orbital Altitude [km]
A controlled de-orbiting of a spacecraft resulting in
an immediate atmospheric re-entry at a predefined
position is only feasible with a high-thrust
propulsion system, at least for the final braking. In
order to confine the impact area in its size to a few
hundred km2, it is necessary to have a sufficiently
steep flight path angle at the atmospheric re-entry.
Simulations predicting the destruction of the
spacecraft during re-entry and the impact area size
have shown, that the Perigee of the transfer orbit
leading to the re-entry has to be 60 km or lower. This
requires to have a flight path angle at the
atmospheric re-entry window, defined here at 120
km, in the range of -1.5° ÷ -2.5°, depending on the
ERS-1&2
m=2400 kg-S/C
1
900
For low orbits below <≅ 615 km no manoeuvre is
necessary to limit the orbital lifetime to 25 years.
Above ≅ 615 km a ∆V of up to 380 m/s for a circular
orbit with 2000 km initial altitude is required for a
de-orbit with an impulsive Hohmann-transfer. A
low-thrust-manoeuvre requires in the range of up to
670 m/s for a very compact spacecraft (Safir-2) with
a 2000 km initial LEO. This big difference in overall
∆V between a continuous low-thrust manoeuvre and
an impulsive manoeuvre, especially for high initial
orbits, led to the conclusion that, even with lowthrust systems like electric propulsion, it is more
effective to subdivide the overall braking manoeuvre
into a number of sub-manoeuvres with braking only
in the Apogee-region of the orbit. This multi-burn
strategy limits the required overall ∆V to a value
close to that of a single-burn Hohmann-type
manoeuvre. Furthermore, the accumulated thruster
operating time is reduced and operation from battery
power in eclipses is avoided for E.P. if periapsis
direction is chosen properly. On the downside, this
choice leads to an increased mission time and the
requirement of cycled thruster operation.
IRS-1C
Munin
800
Figure 3: ∆V-requirement for uncontrolled
disposal to orbits with 25 years remaining lifetime
1000
Safir-2
700
Ref_SC_Lifetime(Defined Lifetime)
Figure 2 shows the remaining orbit life-time of
spacecraft as a function of the Area-to-Mass ratio
and the altitude. The main influencing factor for the
orbital lifetime is the initial orbital altitude. In the
shown range of Area-to-Mass-ratios, the lifetime
varies from a few days to a few months for an initial
orbital altitude of 300 km to several hundred or even
thousand years at 900 km initial altitude. The Areato-mass ratio has a relative strong impact for
compact satellites (Area-to-Mass ≤≈0.005 m2/kg).
For very large and lightweight spacecraft the
importance of this factor decreases. Three horizontal
lines limit the range with 15, 25 and 40 years
maximal orbit life. Spacecraft with an circular orbit
of ≈600 km are in the vicinity of these boundaries.
100
∆V required for disposal to orbit with limited
lifetime 25 years
100
∆V-Requirements
0,014
Area-to-Mass Ratio [m^2/kg]
Figure 2: Orbital lifetime of satellites
Figure 3 shows the required ∆V as a function of the
initial orbital altitude for disposal of the spacecraft to
an orbit with 25 years remaining lifetime. Two
braking strategies are compared. The first strategy
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Deutscher Luft- und Raumfahrtkongress 2002
END-OF-LIFE DE-ORBITING Strategies for Satellites
R. Janovsky et al.
DGLR-JT2002-028
Delta-V [m/s]
initial orbital altitude before the final braking
manoeuvre. Figure 4 shows the required ∆V to
perform a controlled re-entry as a function of the
flight path angle in 120 km and the initial orbital
altitude.
Residual Mass [%]
100
Pathfinder res. Mass share [%]
Residual spacecraft mass after
atmospheric re-entry (impact mass)
Munin res. Mass share [%]
90
Safir-2 res. Mass share [%]
Abrixas res. Mass share [%]
80
IRS-1C res. Mass share [%]
70
60
500
50
450
40
2000 km
Hohmann-Transfer:
Initial circular orbit
30
400
20
10
350
0
300
0
sat_destruction.dia
0,5
1
1,5
2
2,5
3
3,5
4
Flight Path Angle at Re-entry [°]
250
Figure 5: Impacting mass share of reference
spacecraft as function of re-entry flight path
angle
1000 km
850 km
200
650 km
500 km
150
The main results of the destruction prediction are as
follows:
100
-4
-3,5
re-entry conditions
-3
-2,5
-2
-1,5
-1
-0,5
0
Flight Path Angle at 120 km [°]
Figure 4: Effect of initial altitude and re-entry
flight path angle on ∆V (circular orbits)
■ Very small satellites burn up during re-entry for
From Figure 4 it is obvious, that the ∆V-requirement
varies strongly with the initial orbital altitude.
Whereas a satellite with an initial altitude in the
order of 500 km needs only about ∆V500 km=150 m/s,
a satellite in 2000 km requires more than ∆V2000
km=450 m/s.
■ Very heavy satellites survive the re-entry at least
all initial conditions. Therefore no care has to be
taken for this case.
partially in any case. Therefore in this case the
re-entry should be controlled and steep enough
to ensure a well-defined impact area and
location.
■ For medium-sized satellites it depends on the
Regardless of the examined option ‘controlled’ or
‘uncontrolled’ de-orbit, a substantial braking
manoeuvre with up to some hundred m/s ∆V has to
be performed.
initial conditions whether they reach ground in
larger pieces or not.
■ For a steep re-entry it is more likely that larger
pieces reach ground, but in this case it is well
known where they will impact.
■ For a shallow re-entry it is more likely that the
Destruction prediction
satellite breaks up into small pieces, and the
fragment dispersion can become large, but the
fragments are likely to burn up during their way
to ground or loose at least enough mass to
represent no severe hazard when impacting.
As mentioned before, the decision whether a
controlled de-orbit is required or an uncontrolled deorbit is permissible is dependent on the residual risk
for the ground population. This mainly determined
by the number and size of spacecraft parts surviving
the re-entry and impacting on ground. An upper limit
of the so-called “Debris Casualty Area” of about 8
m2 must not be exceeded to permit an uncontrolled
de-orbit. Thus, it is necessary to predict the
destruction of the spacecraft during atmospheric reentry. The destruction of the reference satellites
during their re-entry flight through the Earth
atmosphere has been computed with the SCARAB
software. For each satellite a geometrical model was
constructed, materials were assigned to all geometric
elements, and for selected initial orbits the re-entry
motion and destruction was computed. Figure 5
shows the impacting mass share.
Influence of Atmospheric Drag
A controlled de-orbit manoeuvre requires a stable
satellite trajectory and attitude until completion of
the last de-orbit boost. In the lower atmosphere the
strongest disturbing forces and torques are the
aerodynamic actions. Aerodynamic force and torque
depend on the satellite geometry and on the
atmospheric conditions. During re-entry they vary by
many orders of magnitude. This variation is mainly
due to the linear dependence on the atmospheric
density. The forces and torques also scale with the
size of the satellite.
During re-entry the aerodynamic drag equals the
gravitational attraction along the trajectory usually
twice: the first time, when the satellite leaves its high
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Deutscher Luft- und Raumfahrtkongress 2002
END-OF-LIFE DE-ORBITING Strategies for Satellites
R. Janovsky et al.
DGLR-JT2002-028
altitude orbit and reaches the lower atmosphere; the
second time, when the satellite approaches ground.
The first case occurs typically at an altitude of about
95 km. This means, that the drag can be neglected in
the considerations about the de-orbit control.
The relation between the aerodynamic torque and the
gravity-gradient torque depends strongly on the
aerodynamic moment coefficient of the satellite.
Typically both torques become comparable between
250 km and 300 km altitude. The aerodynamic
torque decreases with the satellite size, but the
necessary thrust to control the attitude increases with
size. The aerodynamic torque depends strongly on
the satellite shape and on the location of the centre of
mass. This implies a lower limit for the altitude, at
which a given satellite (shape, size) can be deorbited in a controlled way with its on-board attitude
control capabilities. This is especially true for ACS
with electrical propulsion thrusters.
If the required attitude cannot be maintained at
perigee, the satellite will leave the aerodynamic
torque dominated region with a superimposed (free)
rotation, which may be difficult to compensate and
may inhibit any subsequent directed de-orbit thrust
manoeuvres. This is even true for satellites which
have an aerodynamically stable attitude and which
will oscillate at perigee about this attitude.
From the aerodynamic point of view it may be
concluded, that a single-burn manoeuvre with the
aim of an immediate re-entry should be performed
above 300 km altitude (This restriction can be
relaxed for very heavy or fast spinning satellites).
Multi-burn manoeuvres or low-thrust spiraling reentries have to consider the torque to be provided by
the onboard ACS to control the attitude against the
aerodynamic torque below 250 km altitude.
Solid
Propulsion
•
•
•
•
•
Simple
Reliable
Low cost
High density
Low structural
index
Hybrid
Propulsion
•
•
•
•
•
Simple
Modulable
Low cost
Reliable
Very high Isp
Electrical
Propulsion
• One thruster per
burn
• Total Impulse fix
• Currently not
qualified for longterm space
application
• Not qualified
• Lack of suitable
oxidiser for longterm mission
• Low thrust
• Complex
• Large manoeuvre
time
• Power
consumption
Table 2: Propulsion system options
The following figures show the specific Impulse and
the thrust range of the different concepts. It has to be
noted that the thrust varies over 10 orders of
magnitude, while the specific impulse varies over 2
orders of magnitude.
Thrust Range
Bipropellant
Monopropellant
Cold Gas
Arcjet
Resistojet
SPT / HCT
Ion
PPT
FEEP
0.000001
0.0001
0.01
Solid
1
100
10000
Thrust [N]
Figure 6: Thrust range of different spacecraft
propulsion systems
Isp Range
Solid
De-Orbit Propulsion System Options
Bipropellant
Propulsion systems investigated in this study are
chemical propulsion systems (including cold gas) as
well as solar-electric propulsion systems. The
following table gives a short overview over principal
characteristics of spacecraft propulsion systems
Advantages
Cold Gas
Mono
Propellant
BiPropellant
(storable)
•
•
•
•
•
•
•
•
•
•
Mono-propellant
Cold Gas
0
50
100
150
200
250
300
350
Isp [s]
Disadvantages
Figure 7: Specific impulse range of chemical
propulsion systems
Simple
• Extremely low Isp
Low system cost • Moderate impulse
capability
Reliable
• Low density
Safe
• High pressure
Wide thrust
• Low Isp
range
• (mostly) toxic
Modulable
fuels
Proven
Wide thrust
• Complex
range
• Costly
Modulable
• Heavy
Proven
• Toxic
Isp Range
Arcjet
Resistojet
SPT / HCT
Ion
PPT
FEEP
0
1000
2000
3000
4000
10.000
5000
Isp [s]
Figure 8: Specific impulse range of electrical
propulsion systems
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Deutscher Luft- und Raumfahrtkongress 2002
END-OF-LIFE DE-ORBITING Strategies for Satellites
R. Janovsky et al.
DGLR-JT2002-028
These different propulsions options have been
considered for all reference spacecraft an assessed
for their suitability to perform the EoL-de-orbit
manoeuvre.
P/L MGSE
MLI
Radiator
Heate
r
Rx
Digital
P rocessor
Memory
Tx
Analogue
IF Electr.
Power
HP Cmd
Decoder
BCU
Assuming that the de-orbit function is implemented
in future satellite designs, two baseline concepts
were considered:
Telem etry /
Comm and
BUS DHS
Data Handling System
T
Sensors
Design of De-orbit function
Control/Data I/F
PWR
HKP
BUS
Subsystem s
IF Electr.
& Control
Solar
G enerator
(SG)
B attery
Package
(BP )
Reg. P ower
Power
Conditioning
Distribution
(BUS PDCU)
Power
Electronic Power
Subsystem (EPS)
■ Integrated de-orbit function: The de-orbit
Launcher I/F
function is integral part of the Attitude & Orbit
Control subsystem.
System EGSE
H eat
Pipes
Therm al
Control
Subsystem
(TCS)
S -B and
A ntennas
Mechanical
Therm al I/F
Payload
Power I/F
Structure &
Mechanisms
Attitude (Orbit)
C trl. Subsystem
(AOCS)
IF & Ctrl.
DO F
CMD
O CS
Sensors
Deployable
s
Actuators
Actuators
GPS
DOPS
S/C MGSE
Legend:
■ Additional De-orbit Control Subsystem: An
P ower
extra subsystem is dedicated to the de-orbit
function.
An add-on de-orbit module including own control
and power is recommended in case of existing
satellite design without de-orbit function. Figure 9
and Figure 10 show the generic design-concepts. An
overview of all OCS options treated during the study
is summarised in Table 3. For each OCS option two
thrust levels and two ∆V levels were applied
resulting in a total of more than 100 de-orbit variants
analysed.
DOF = De-Orbit-Function
Interfaces
DOPS = De-O rbit-Propulsion-System
Ctrl
RF Link
TM Data
Env.
Output
E nv.
Input
/D e-Orbit-Func tion_v2.v s d
Figure 9: Integrated DeOrbit Function Design
P/L MGSE
Radiator
T
Sensors
H eate
r
BCU
IF Electr.
& Control
BP
PDCU
Control/Data I/F
Telemetry /
Comm and
BUS DHS
Data Handling System
Rx
Digital
Processor
M emory
Tx
Analogue
IF Electr.
Power
HP Cmd
Decoder
PW R
HKP
BUS
Subsystems
SG
AOC S
EPS
Structure &
Mechanisms
CM D
Deployable
Actuators
Power
IF & Ctrl.
DeO rbit C trl. Subsystem
(D OCS)
Power
IF & Ctrl.
O CS
Sens.
Act.
G PS
System EGSE
MLI
Heat
P ipes
Therm al
Control
Subsystem
(TCS)
S-Band
Antennas
Mechanical
Thermal I/F
Payload
Pow er I/F
OCS
Sens.
Act.
G PS
Launcher I/F
S/C MGSE
Legend:
Power
Interfaces
Ctrl
TM Data
R F Link
Env.
Ou tpu t
En v.
Inpu t
/De-Or bit-F unc tion-add- on_v1.v sd
Figure 10:Additional DeOrbit Control Subsystem
Design
Pathfinder
Munin
Cold Gas
Cold Gas
Solid Propellant Solid Propellant
Mono Prop.
Mono Prop.
Bi Propellant
Bi Propellant
N/A
N/A
N/A
N/A
*for uncontrolled de-orbit only
Table 3: OCS-combinations
Safir
Cold Gas
Solid Propellant
Mono Prop.
Bi Propellant
PPT*
N/A
Abrixas
Cold Gas
Solid Propellant
Mono Prop.
Bi Propellant
SPT*
EHT*
Two manoeuvres have been defined, to investigate
the suitability of the different propulsion system
options:
■ Manoeuvre 1:
■ Manoeuvre 2:
IRS
Cold Gas
Solid Propellant
Mono Prop.
Bi Propellant
SPT*
Ion*
2420-kg-Sat
Cold Gas
Solid Propellant
Mono Prop.
Bi Propellant
SPT*
Ion*
Circular orbit
Manoeuvre ∆V
400 km
100 m/s
780 km
200 m/s
Table 4: Maximum initial orbital altitudes for
controlled de-orbit (perigee 60 km)
∆V=100 m/s
∆V=200 m/s.
With these manoeuvres, the following de-orbiting
tasks could be performed (controlled de-orbit Table 4, uncontrolled de-orbit Table 5):
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Deutscher Luft- und Raumfahrtkongress 2002
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R. Janovsky et al.
DGLR-JT2002-028
Spacecraft
∆V= 100 m/s ∆V= 200 m/s
Pathfinder
980 km
1370 km
Munin
910 km
1290 km
Safir-2
870 km
1240 km
Abrixas
910 km
1260 km
IRS-1C
910 km
1300 km
2420 kg-spacecraft 870 km
1250 km
strategies, concepts and propulsion systems has
been defined. Using a standardised quantitative
evaluation scheme, an abstract value, the so called
System Cost Figure (SCF) is calculated. The SCF is
the sum of nine terms as follows:
9
SCF =
■ Spin stabilisation as basic strategy for the 'fast'
torques
p
i
i =1
Two generic attitude control strategies have been
identified for the de-orbit task:
disturbing
i
The terms Ai are the evaluation terms, taking into
account e.g. the increase of cost, mass and volume
by implementing the de-orbit function as well as
other parameters like reliability, additional
electrical power required and complexity. A lower
SCF corresponds thereby to a better performance.
The SCF should not be mistaken to have a direct
monetary equivalent, i.e. that the cost of the de-orbit
function is directly related to the SCF figure. The
SCF shows the relative influence of the individual
assessment criterion. In this study, cost, reliability,
mass and volume were selected to have the highest
priority and thus the highest importance in the
selection of the de-orbiting concepts. The quantities
C1...C9 are weighting factors to adjust the terms for
the selection procedure with regard to their
importance.
Table 5: Maximum initial orbital altitudes for
disposal to orbit with 25 years limited lifetime
cases with high
accelerations
∑C A
and
■ Coarse 3-axes stabilisation for the 'slow' cases
and in case of special requirements like Sun
pointing or structural constraints
Spin stabilisation is in general feasible for all
satellite classes, but will be especially chosen in
cases which would need additional AOCS
components. The satellite is spun along the de-orbit
thrust vector axis, which is aligned parallel to the
velocity vector in the apogee (orbit tangent). It was
assumed that in general the de-orbit AOCS is not
capable to align the spin axis permanently parallel
to the velocity vector during the manoeuvre (some
time before and after apogee). Therefore the burntime of a single manoeuvre shall be limited to about
4,5% of the orbit (+/- 8,1°) corresponding to 1%
effective thrust losses. The basic ACS configuration
for spin stabilisation will consist of Sun sensor or
horizon sensors, magnetometer, and magnetic
torquers. 3-axes stabilisation will generally be the
choice in case of low torques and angular
accelerations (i.e. electric propulsion) or in cases,
where spin stabilisation is not feasible (i.e. design
constraints). For bigger satellites an appropriate 3axes ACS will be already existing in the most cases.
The single burn time can be exceeded to 20% of the
orbit. The basic ACS configuration for 3-axes
stabilisation is strongly dependent on the
disturbing torques and accelerations. Beside Sun
sensor or horizon sensors and magnetometer, it may
consist of magnetic torquers and/or momentum
wheels and/or cold gas thrusters.
Comparative Assessment for
Reference Satellites
For the comparative assessment of the de-orbiting
concepts, the two defined manoeuvres generating a
total ∆V of 100m/s respectively 200m/s were used.
The results of the comparative assessment for two
of the selected reference spacecrafts are presented
in the following paragraphs.
Pathfinder
Figure 11: Pathfinder
The Pathfinder-S/C with a nominal mass of 1.6 kg
is the smallest spacecraft considered in this study.
Pathfinder is presently in a conceptual design phase
and intended to investigate the magnetic field of the
Earth. For this purpose, a number of 100-200
satellites shall be flown. Pathfinder has a cold gas
system used for a spin-up to stabilise the spacecraft.
Due to its small size, the implementation of a deorbit function has a tremendous impact on the S/Cdesign, regardless what type of components are
selected. This is mainly caused by the fact, that not
all necessary components for a de-orbit can be
miniaturised and adapted to the small size of the
S/C. Thus the additional mass and volume required
Evaluation Criteria
Mass and volume models have been established to
determine the impact of the various de-orbiting
concept options on the design of the spacecraft. A
set of criteria to evaluate different de-orbiting
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2500
The IRS-1C S/C belongs to the medium-sized
spacecraft and it is designed for Earth observation
purposes. It has a nominal mass of 1250 kg without
de-orbit device. The spacecraft is 3-axes stabilised
and uses a hydrazine OCS. The multi-burn strategy
allows to perform a de-orbit manoeuvre with a
thrust level of less than 10-50 Newton for the
“chemical” propulsion systems. Here electric
propulsion is also an option, but the available power
level is still low. Therefore the manoeuvre has to be
subdivided into some 1000 burns in the apogeeregion, resulting in overall manoeuvre duration of
20 days or more. The impact of the implementation
of a de-orbit function on the IRS-1C is shown in
Figure 14.
The SCFs of the de-orbit options for the IRS-1C
spacecraft are considerably lower than the SCFs for
the Pathfinder spacecraft. It can be deduced from
this fact that the relative impact of the addition of a
de-orbit capability is more profound on the
Pathfinder spacecraft. However, conclusions that
the addition of the de-orbit function is e.g. more
costly or heavier in absolute terms on the Pathfinder
spacecraft as compared to the IRS-1C spacecraft are
not permissible and not valid.
Pathfinder System Cost Figure
De-Orbit Task ∆V=100 m/s,
Single Burn
2000
Capability
Complexity
Cas. Area
Man. Time
Power
Volume
Mass
Reliability
Cost
1500
1000
500
System Cost Figure [-]
System Cost Figure [-]
for the implementation of the de-orbit function are
of the same order of magnitude like the original
S/C. Also the cost linked with the implementation
of a propulsive capability of 100 m/s are roughly of
the same order of magnitude as the original S/C
without de-orbit function, having thus a significant
impact on the overall SCF.
The original spacecraft is designed for an average
electric power of 1 W with a peak power of 5 W.
This is by far too low for the operation of any
electric propulsion system to be used for the deorbiting. Thus only the cold gas system, a solid
propellant system, a mono-propellant system and a
bi-propellant system were included in the
assessment. The impact of the implementation of a
de-orbit function on the Pathfinder is shown
subsequently, showing individual contributions to
the SCF.
0
1:Cold Gas
2: Solid
Propellant
3: MonoPropellant
4: BiPropellant
Figure 12: SCF for Pathfinder-De-Orbit options
IRS-1C System Cost Figure: DeOrbit Task ∆V=100 m/s, Single
Burn
140
120
100
Capability
Complexity
Cas. Area
Man. Time
Power
Volume
Mass
Reliability
Cost
80
60
The solid propulsion system is by far the best
choice with regard to the overall SCF. The
numerical advantage in the overall SCF is in the
order of 40%, compared to the second choice,
which is the mono-propellant system. This result is
linked to the additional cost of the de-orbiting
system, where the solid propellant system is the
cheapest. Also with regard to additional mass and
volume, the solid propellant system is the best
choice, followed by the mono-propellant system
and the cold gas system. Mass and volume increase
only by less than 1 kg respectively 2.5 ltrs, whereas
it is more than +2.7kg respectively 4ltrs for the
mono-propellant system as second best choice.
It is important to note that the assumptions for a
solid propulsion system are based on published data
of small solid motor prototypes. Current
commercially available solid motors have too high
thrust levels and too short burn times, leading to
excessive acceleration levels.
40
20
0
1:Cold Gas
2: Solid
Prop.
3: Mono-P.
4: Bi-Prop.
5: Arcjet
6: Ion
Figure 14: SCF for IRS-1C-De-Orbit options
For IRS-1C the solid propellant propulsion system
as well as the mono-propellant system are
performing comparable with regard to the overall
SCF. The solid propellant system has some
advantages with regard to cost, mass and volume,
the mono-propellant system with regard to
operational flexibility. The bi-propellant system is
only a few percent behind. The two E.P-options are
ranked on 4th and 5th place with some distance. The
cold gas system is clearly on the last rank due to its
mass- and volume-penalty, which also has an
impact on the overall cost (increasing launch cost).
The implementation of a de-orbit function into the
IRS-1C-S/C would require additional cost of about
2÷3M€ in case of a solid propellant system, which
is the cheapest solution. This is in the order of 2%3% of the overall S/C-cost without de-orbit
function. The mass impact is in the order of 60kg or
5% in case of braking manoeuvre with a total
∆V=100m/s.
Due to the fact that the available power level is to
be considered moderate, time needed for de-
IRS-1C
Figure 13: IRS-1C
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R. Janovsky et al.
DGLR-JT2002-028
orbiting using E.P. becomes relatively long. Fuel
consumption is relatively low, yet, E. P. hardware
masses kill a considerable fraction of the fuel saving
advantage. E.P. turns out not to be the solution of
choice in this configuration. If in contrast E. P. is
already installed on the S/C for the nominal mission
and there is no problem to use the S/S for deorbiting as well (requirement for thrust vector
Satellite
Propulsion System
Cold
Gas
Pathfinder
Munin
Safir
Abrixas
IRS-1c
2420kg-Sat
directions while respecting compatibility with solar
generator orientation and performance), E.P. will
perform better in the comparative ranking.
Solid
Prop.
Mono
Prop.
BiProp.
Arcjet
Ion
3
1
2
4
N/A
N/A
3
1
2
4
N/A
N/A
4
1
2
3
N/A
N/A
6
1
2
3
4
5
6
1
1
3
4
5
6
1
2
2
4
5
Table 6: Ranking of de-orbiting options for all spacecraft
exclusively with electric propulsion, since at least
the final burn requires to generate a significant ∆V
outside the atmosphere, requiring a sufficient thrust
level not available with electric propulsion. In case
of a controlled de-orbit, a combination of a highthrust system with electric propulsion or a pure
chemical propulsion system is needed.
Although the absolute effort to install a EOL-deorbit function into a spacecraft increases with the
size of the spacecraft, the relative impact is most for
the small vehicles. Whereas a spacecraft of the
Pathfinder-class has a SCF in the order of
SCFPathfinder, best=500, it decreases by a factor of 10
for the >2000kg class. Thus it is more easy to
implement the EoL-de-orbit function into the larger
spacecraft, where the overall budgets are increased
only by a few percent.
Synthesis
The results of the ranking for all investigated
spacecraft and propulsion system options is shown
in Table 6.As a general result, cold gas propulsion
systems were found to be limited in their
applicability as de-orbit propulsion system to
relatively small satellites (mSat ≤≈ 500 kg), due to
the severe mass impact. Electric propulsion systems
are most useful in big spacecraft, (mSat ≥≈ 1500 kg),
having either electric propulsion for other reasons
(e.g. station keeping) or having installed substantial
electric power under the assumption of uncontrolled
de-orbit. The solid propellant system is for almost
all spacecraft the best solution to perform an end-of
life de-orbit manoeuvre, followed by the monopropellant system. This ranking is relatively
insensitive to the size of the manoeuvre. Electric
propulsion can be used as de-orbit option only if the
spacecraft exceeds a minimum size. For smaller
spacecraft, the additional electric power to be
installed makes the use of the electric propulsion
not reasonable. For larger spacecraft and braking
manoeuvres, the Isp of the propulsion system is
getting more and more important. Thus, electric
propulsion become more and more competitive.
The reason for the worse ranking of E.P. is, that in
any case considered here it was needed to install
additional electric power to keep in case of electric
propulsion the manoeuvre time below the specified
one year. This disadvantage with regard to mass,
cost, volume, complexity could not be compensate,
resulting in a 4th rank for electric propulsion at the
best. If it is permissible to extend the overall
manoeuvre duration, the electric propulsion
becomes more and more attractive, especially for
larger manoeuvres as required for an uncontrolled
de-orbit from very high initial altitudes. A
controlled de-orbit can not be performed
Conclusion
Propulsive methods for end of life de-orbiting of
spacecraft passing through LEO regions have been
assessed and classified. Satellites having an orbit
above ≅ 615 km altitude have a natural lifetime of
more than 25 years and need an active de-orbit
manoeuvre. The transfer to an elliptical orbit with a
limited lifetime of 15 to 40 years has emerged as
the optimal solution for satellites where an
uncontrolled re-entry is admissible. This strategy
leads to a minimal ∆V requirement. In cases that a
controlled re-entry is necessary, a flight path angle
between -1.5° and -2.5° has to be achieved at the reentry window at 120 km altitude to limit the
ground-impact footprint.
The addition of a de-orbit function on a spacecraft
has in most cases a significant effect on satellite
design due to the fact that the ∆V requirements for a
de-orbit can rise up to 450 m/s and above. The
impact on small spacecrafts is thereby even greater
as the nominal mission usually does not require an
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OCS and propellant hardware dry mass has a
minimum value. It is found that the spacecraft mass
of micro- and nano-satellites can double due to the
addition of the de-orbit function. Miniaturization
efforts and ongoing developments in this domain
could improve the situation.
Solid propulsion appears as an attractive solution in
many cases due to its compactness and a relative
small mass increase. However it has to be
mentioned that adequate solid motors with respect
to thrust level and burn duration exist solely as
prototypes nowadays.
Monopropellant thrusters seem to be a good
alternative in most cases, having the advantage of
being readily available. Bipropellant thrusters seem
to be competitive only on heavy satellites.
Cold gas systems are tremendously impaired by
their low Isp. A use for a de-orbit function seems
therefore only advantageous for small satellites and
low orbits with extremely small ∆V requirements.
Electric propulsion was not identified as an optimal
solution for any of the chosen reference satellites.
This is due to the fact that none of the spacecrafts
was equipped with E.P. for nominal mission life and
that available electric power was low. It can be
expected that the outcome is considerably different
for sufficiently heavy satellites with a sufficiently
large power/mass ratio and a relative high initial
orbit.
This study was intended to highlight advantages and
drawbacks of different EoL-de-orbiting strategies
and to give a general indication on suitability of
different technical options for de-orbit tasks.
However the selection of a propulsion system will
have to be made in a case by case study, taking into
account the total ∆V requirement of the satellite,
both for the nominal mission and the de-orbit
function, synergetic effects and possibly other (e.g.
non-technical) boundary conditions.
Acknowledgements
The work presented in this paper was performed by
OHB System, DLR and HTG within the General
Study Programme of ESA, ESA-contract
15316/01/NL/CK.
References
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Release 1.0, April 7, 1999
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Proceedings of the first European Conference
on space debris, ESA SD-01, pp 607-615, 1993
[3] NN: The Orbital Debris Quarterly News – A
Decade of Growth, NASA Johnson Space
Center, Volume 5, Issue 4
10