Designing a space debris removal mission targeting a Cosmos 3M
rocket body in LEO using a chemical engine and an elechtrodynamic
tether to perform the de-orbiting.
Abstract
During the past years, several studies have
been conducted to depict an image of the
current state and future evolution of orbital
regions, showing that space debris mitigation
is necessary. Amongst others, this led the
IADC to develop a series of (passive)
mitigation guidelines adopted in a 2007 UN
resolution. These studies which evaluate longterm debris evolution have indicated that the
debris density has reached such a high level,
that there will be an ongoing increase of the
number of debris objects, primarily driven by
collision activity. This effect will manifest
itself mainly in the Low Earth Orbit (LEO)
region, due to a combination of high spatial
densities, high relative velocities and a large
number of object crossings[1]. An analysis of
potential collisions enabled the selection of a
limited number of large, intact objects
imposing the greatest risk for future
instability of the LEO environment; in fact,
on the long term, the large >10 cm objects
will play a more critical role. As they contain
such a majority of the mass, they can form the
source of giant amounts of new, smaller,
debris with a cascade effect[2]. The passive
mitigation measures currently used do not
suffice as an insurance of a stable debris
environment for the next years endangering
the future of the space mission not only in
LEO and an active removal mission to clean
became clearly a necessity.
One of the concepts for an active debris
removal system (ADRS) consists on using a
satellite or an adapted launch vehicle upper
stage to enable the collection and de-orbiting
of debris elements. The ADRS should be
equipped with various subsystems, including
a small propulsion system for approaching the
debris through a far-guidance orbital
maneuver, a tethered space micro-tug (SMT)
for performing the close-range rendezvous
and a docking system for establishing
physical connection with the debris
object[3][4]. Approaching the space debris has
to be performed using a chemical engine,
because it can assure a quick rendezvous with
a high-velocity object such as a space debris
in LEO. After the debris is grabbed, the object
has to be de-orbited and burn into the
atmosphere, as in LEO there is no such option
like a cemetery orbit. This could be performed
either by the same chemical engine or by an
elechtrodynamic tether[5]. Factors that should
be considered in choosing the de-orbiting
approach are the mass of the system (included
propellant tanks or additional batteries), time
of de-orbiting, control authority over the
whole system (satellite + space debris) to
avoid collision during the final stage of the
mission and the efficiency of the system.
In this project, we will use the AGI STK
software package to study the difference
between the two configurations and
eventually reach a decision as to which
approach
is
recommended.
Starting Literature:
[1] Valery I. Trushlyakov, Matteo Emanuelli, Alexander Ronse, Claudio Tintori; A Space Debris Removal Mission using the orbital
stage of launchers, Heinelin “Flight Into the Future” 2011 contest, Moscow 2011
[2] D.J. Kessler & B.G. Cour-Palais; Collision Frequency of Artificial Satellites: The Creation of a Debris Belt, Journal of
Geophysical Research, Vol 83, June 1978, Pages 2637- 2646
[3] Trushlyakov V.,Jakovlev M., Shatrov J, Shalay V.V, DeLuca L.T, Galfetti L., Active de-orbiting system of SLC upper stages and
spacecraft based on hybrid gas rocket engines, Book of 3-th European Conference on Space Science, Versailles, France, 6-9 July,
2009
[4] Trushlyakov V, Shalay V., Shatrov J., Jakovlev M., Kostantino A., Аctive de-orbiting onboard system from LEO of upper stages
of launchers, Proc. “5th European Conference on Space Debris”, Darmstadt, Germany, 30 March – 2 April 2009, (ESA SP-672, July
2009)
[5] Claudio Bombardelli, Javier Herrera-Montojo, Ander Iturri-Torrea and Jesus Peláez Electrodynamic tethers for space debris
removal, 2010,Technical University of Madrid, Spain
Aim & Steps
Using a satellite or an adapted launch vehicle upper stage to enable the collection and de-orbiting of
debris elements.
 Investigation on objects that can be tracked and disposed of (size, mass, altitude).
 Modeling the satellite and space debris.
 Mission phases: far approach(1), close approach(2), de-orbiting(3).
1. Composition of the system proposed (far approach): satellite (+ tether system + micro tug).
2. Composition of system proposed (close approach): satellite + tether system (20 km) + micro tug.
3. Composition of system proposed (de-orbiting): satellite + tether system (20 km) + micro tug +
space debris.

How to perform de-orbiting after grabbing?
 Chemical engine: one or more burns to reduce altitude and provide de-orbiting of the whole
system.
 Elechtrodynamic tether: using the interaction of the magnetic field with a current that flows in
the tether, to generate a force to pull down the whole system through the atmosphere to the
ground.

Compare the result.
Accomplished Milestones





Investigation on object that can be tracked and dispose (size, mass, altitude).
Modeling the satellite and space debris
Investigation on object that can be tracked and dispose (size, mass, altitude).
Definition of launcher and launch site
Initial simulation of the launch
Definition of the orbit
The identification of the orbital region where the space debris situation is more critical is the
starting point of this research.
LEO or GEO?
89%of the ~950 operational satellites are either in a low earth orbit (LEO, 300-2000 km altitude) or
a geosynchronous orbit (GEO, ~36000 km altitude). Hence, these two regions form the first focus
of our selection.
Figure 1: Distribution of space debris in different orbits
Both have specific characteristics regarding the presence and evolution of space activities, but in the
LEO region, satellites and debris elements are quite widely scattered in terms of altitude, inclination
and ascending node. This, in combination with the fact that orbital speeds are considerably higher
than in GEO, makes both the amount of crossings and the relative velocities of the bodies during
these crossings very high. The wide and random distribution of objects also implies that a system of
graveyard orbits (as in the GEO case) is not practical. Another critical issue is that manned spacemissions are performed at (low) LEO altitudes, making it essential that the risk of collision is
minimized to the greatest possible extent. So, as described in [5], the combination of a higher
debris concentration, a large number of crossings and high relative velocities in the LEO region
may lead to an exponential growth of debris objects by a future cascade of collisions.
Orbital parameters
Once the LEO region is defined as a primary target of our investigation, the most critical orbital
parameters must be identified to choose a proper target for the mission.
Studies were performed for predicting the probability of collision in the next centuries using
NASA’s LEGEND model, based on the past and current debris environment. These studies form the
basis of further orbital selection [1][4]. The collision probability in the next 50 years is higher at the
altitude bands containing the highest fragments concentration caused by the catastrophic events of
Fengyun 1c at ~850km and Iridium-Cosmos at ~800km. The critical altitude band is extended to
800-1000 km.
Figure 2- Normalized distribution of predicted catastrophic collisions as a function of altitude and spatial density
(1). (Distributions, for objects 10 cm and larger, at the end of 2005 and 2205)
Higher inclinations (60°-110°) are much more crowded as a direct result of the high number of past
and present satellites which use these zones to fulfill their mission goals. The peak between 90° and
100° is due once again to the high amount of fragments of Fengyun 1c. Thus, efforts for actively
remove debris should focus on objects in those orbits.
Figure 3 - Amount of detected LEO objects per inclination band for 800-1000km altitudes. Results are based on
SSN data of October 2010 (using (2) and (3)) and inclination distribution for objects involved in future LEO
collisions (4).
Unlike the semi-major axis and inclination, no particular trend can be seen in the right ascension of
the ascending node (RAAN) of the current debris environment and future collisions. This is due to
the oblateness of the Earth, which makes space debris RAAN a permanently evolving parameter.
Figure 2: Evolution of debris cloud from Iridium 33 and Cosmos 2251 satellite collision in 2009 (5).
Image (a) shows the situation at the time of the collision, (b) depicts the orbits of the various debris
objects 6 months later.
Figure 4 - Amount of detected LEO objects a function of eccentricity. Results are based on SSN data of October
2010 (2) (3).
Definition of the target
Using the USSTRATCOM TLE database, which contains the Keplerian elements of all detectable
debris objects, a list was made of all candidate disposable debris objects. The data was filtered
based on object size and type as well as the most populated orbital regions. When the rocket bodies
are counted per type, the distribution in Figure 5 is reached.
Figure 5 - Types of rocket bodies in the critical region
As it can be seen in the Figure 5, the Russian Kosmos 3M rocket bodies are the perfect candidate
for our research due to the large number of bodies in orbit. Figure 6 illustrated the spatial
distribution of Kosmos 3M 2nd stage rocket bodies around the globe. The image was taken using the
Google Earth program combined with information from the U.S. Space Track catalog and the UCS
Satellite Database
Figure 6 - Spatial distribution of Kosmos 3M rocket body in LEO
The STK software and Space Track database were used to define the low Earth orbit environment
according to spatial distribution of Kosmos 3M rocket bodies, as outlined in Figure 7.
Figure 7 - Spatial distribution and orbits of Kosmos 3M rocket body in LEO
Orbits of 156 rocket bodies are illustrated in Figure 7 with boundaries set at a perigee of 800 km
and 1000 km. Only 15 rocket bodies are on an inclination between 0° and 80° (light blue on the
figure). On the other hand, for the band of interest (80°-100°), 141 rocket bodies were listed (blue
on the figure). The majority of these objects are at an inclination of around 80°.
Of all the bodies available, one was arbitrarily chosen as the target of the mission. This body is only
chosen to define a representative mission and may change based on future studies. According to the
US Space Track catalog, the rocket body is classified as SL-8 R/B 32053. The orbital parameters of
this object are summarized in Figure 8 and Figure 9.
Figure 8 - Orbital parameters of the SL-8 R/B 32053 object
Figure 9 - Evolution of semi-major axis of the SL-8 R/B 32053 object
Kosmos 3M
Kosmos 3M is a two-stage launch vehicle. It was developed from the Soviet Union's R-14
intermediate range ballistic missile. The rocket, developed by KB Yuzhnoe and manufactured by
PO Polyot in Omsk, consists of an R-14 first stage topped by a purpose-built "S3" restartable
second stage, developed from scratch.
Kosmos 3M is 2.4 meters in diameter, 32.4 meters tall, and weighs 109 tons at liftoff. Its first stage
is powered by a 151.5 ton sea level thrust (178 ton vacuum thrust) Energomash RD-216 engine
composed of two dual-thrust chamber RD-215 engines. Four graphite vanes extend into the exhaust
of the four fixed thrust chambers to provide steering.
The rocket uses an upgraded second stage with an RD-219 (11D49) fixed, restartable main engine
and a secondary low-thrust on-orbit propulsion system consisting of four steerable thrusters. This
feature allows the vehicle to maneuver while releasing multiple satellites during a single mission.
The four steering thrusters are fed from two side tanks mounted on the exterior of the stage.
Table 1 - Vehicle components (6)
Diameter (m)
Length (m)
Propellant Mass (tons)
Total Mass (tons)
Engine
Fuel
Oxidizer
Thrust
(SL tons)
Thrust (Vac tons)
ISP (SL sec)
ISP (Vac sec)
Burn Time (sec)
No. Engines
Stage 1
2.4 m
21.5 m
81.9 t
87.2 t
RD-216
UDMH
Nitric Acid/
27% N2O4
151.5 t
Stage 2
2.4 m
6.5 m
18.7 t
20.14 t
RD-219
UDMH
N2O4
PL Fairing
2.4 m
4.5 m
178 t
248 s
291.3 s
131 s
1
(2x2chmbr
RD-215s)
16 t
t
303s+176s
350s+350 s
1 fixed main
+ 4 steering
s
s
0.5 t
Figure 10- Kosmos 3M schematics (6)
The first R-14-derived launch vehicles (Kosmos 1 or 65S3) flew from Baikonur Area
41 Pad 15 beginning in 1964. The upgraded developmental Kosmos 3 (11K65)
version began launching from the same site in 1966. One year later, the operational
Kosmos 3M (11K65M) variant began flying from Plesetsk Northern Cosmosdrome.
Baikonur developmental launches ended in 1968, but Kapustin Yar began hosting
Kosmos 3M launches in 1973. By 1977, Kosmos 3M had entirely replaced the
smaller R-12-based Kosmos 2 launch vehicle.
Kosmos 3M launched Tselina-O electronic intelligence satellites, Strela 1M, 2M and
Tsyklon military navigation and communications vehicles, Taifun-1 and 2 radar
calibration spacecraft, anti-satellite weapon targets, and Tsikada navigation satellites
into orbit. Orbital and suborbital R-14 launches boosted BOR space plane tests during
the 1980s for the Soviet space shuttle development effort. Other suborbital R-14
missions were performed for the "Vertikal" scientific research program during the
1970s and 1980s. The United Start consortium has marketed Kosmos 3M for
commercial missions since the 1990s.
Kosmos 3M ranks third among world space launchers with nearly 450 orbital
attempts, trailing only R-7 and Thor/Delta. From a peak of 29 launches in 1977, the
Kosmos 3M flight rate has declined to an average of 2-3 per year since the mid1990s. Launch vehicle production in Omsk was stopped in 1995, but production
capability was retained. Stockpiled rockets have reportedly performed subsequent
launches. Kosmos 3M is the only launch vehicle to have operated from all three of
the former Soviet launch sites at Baikonur, Plesetsk, and Kapustin Yar. Kosmos 3M
is integrated with its payload in a processing building and transported horizontally to
its launch pad on a railroad carrier. The rocket and its carrier are erected at the pad
about two days before launch. The Kosmos 3M vehicle was last launched in 2010,
with a total of about 442 launches. (wikipedia)
Figure 11: Installation of Kosmos 3M on the launch pad
Figure 12: Kosmos 3M in Omsk
Kosmos 3M’s de-orbiting removal justification
Considering the size and mass of a Kosmos 3M 2nd stage and the mean value of the orbital
parameters seen previously, the orbit of a typical Kosmos 3M 2nd stage was modeled in STK, as
illustrated in Figure 13.
Altitude: 900km; Inclination: 80°; Propagator: J2M
Figure 13: STK model of a typical Kosmos 3M second stage orbit
A lifetime analysis was carried out (Figure 14) and the result was that Kosmos 3M can’t decay
automatically (limit at 64km) before 25 years, which is a requirement by UN COPUS space debris
regulations. Actually, its orbital altitude only decays to 70km in 100 years, limit imposed for the
STK simulation.
Figure 14 - Limit of decay for the object
Definition of launcher and launch site
After the definition of the target, the next step is the definition of the launch site and launcher. The
choice depends on the target orbit and the availability of a suitable launch vehicle.
According to these boundaries, the potential launch sites were shortlisted among Vandenberg (USA
- 34°43′57″N 120°34′05″W), Kourou (French Guyana - 5.305°N 52.834°W) and Plesetsk (Russian
Federation - 62°57′35″N, 40°41′2″E). The reason why the Baikonur Cosmodorme was not
considered as a possible choice, was the possibility that Russia will abandon all facilities for
Baikonur and launch all satellites from Plesetsk in the next 5 years, based on news reports.
Another important feature of the desired launch vehicle should be the final stage restarting capacity.
Our approach is integrating all de-orbiting subsystems onto the upper stage and using the
upperstage for propulsion purposes, which requires restartability. Moreover, several maneuvers will
be required to reach the target orbit, which will also depend on restartability.
Another parameter that affects the choice of launcher, is the amount of propellant capacity of the
launcher’s upper stage. The upperstage is required to inject a satellite into orbit and also move our
system to the target orbit and back to a LEO orbit for the application of the tether system. As the
amount of propellant used in this process will depend on the upper stage itself, the analysis will be
done with one upper stage. If the upper stage was unable to perform our mission, a secondary upper
stage with larger propellant capacity will be used.
Based on all these parameters, a list of active launch vehicles was created, as outlined in Table 1.
This list summarizes the polar launch capability, upper stage properties and the availability of a
model in STK. Considering all these required characteristics, the Soyuz vehicle with the fregat
upper stage was chosen as the prime launch vehicle. The Proton -M launch vehicle with the BreezeM upper stage will be considered as a secondary alternative system. The Plesetsk cosmodorme was
chosen as the launch site for both vehicles. The Plesetsk cosmodrome is a military site in the NorthWest Russia at some 800 km from Moscow.
Figure 15 - Plesetsk Cosmodrome (6)
Table 2 - Launch vehicles and launch sites (6)
Upper Stage
Launch
Vehicle
Country
Dnepr-1
Ukraine
Zenit
PEO
characteristics
Payload
(kg)
Launch
Site
Dimensions
(m × m)
Dry
Mass
(t)
Propellant
Mass (t)
Total
Lift-off
mass (t)
Restar
table
STK
model?
1,91
4,26
N
Y
-
89,9
N
Y
36
40
N
N
35
39
N
N
Solid
-
0,98
N
Y
0,126
Solid
0,771
0,897
N
Y
Propulsion
RD-869
(UDMH-N2O4)
(KeroseneLO2)
YF-22 B
(UDMH-N2O4)
YF-22 (UDMHN2O4)
800 km, I = 87.3
400
Baikonur
1*3
2,35
Russia
200 km, I = 99
11380
Baikonur
10.4 * 3.9
8,3
CZ-2D
China
200/400 km, I =
90
2750/1175
JSLC
10.4 * 3.35
4
CZ-2C
China
600 km, I = 90
800
JLSC
7.5 * 3.35
4
Taurus-XL
US
400 km, I = 98
880/1050
VAFB
2.1 * 0.97
Minotaur
US
740 km, I = 98.6
335
VAFB
1.34 * 0.97
800 km, I = 98.2
580600/740
VAFB
1.34 * 0.98
Solid
0,77
0,893
N
Y
5,444
5,98
N
Y
9,7
11
Y
Y
12,5
14,7
Y
N
2
2,92
Y
Y
Taurus
US
CZ-2E/ETS
China
1000km, I = 86
4930
XLSC
2.936 * 1.7
0,541
SPTM-17
(solid)
Ariane 5
EU
800 km, i = 98.6
9500
ELA3
3.356 * 3.936
1,25
AESTUS
(MMH-N2O4)
GSLV
India
407 km, I = 51.6
5000
SHAR
8.7 * 2.9
2,2
PSLV
India
800 km, I = 99.1
1200
SHAR
2.65 * 1.34
0,92
Delta-4H
US
500 km, I = 90
21700
VAFB
13.7 * 5.13
3,49
2.850/
3.490
Delta-4M
US
500 km, I = 90
7350-11700
VAFB
12.2 * 4.07 or
13.7 * 5.13
Delta-2
US
833 km, I = 98.7
1591-3186
VAFB
5.88 * 2.44
0,95
ATLAS 2,
2A
US
185 km, I = 90
5510-6170
VAFB
10.06 * 3.05
1,84
ATLAS 5
US
189 km, I = 90
9050-10750
VAFB
12.68 * 3.05
1.914/
2.106
CZ-3
China
200 km, I = 90
3000
XLSC
7.48 * 2.25
2
CZ-4
China
900 km, I = 99
1650-2800
TSLC
1.92 * 2.9
1
Soyuz-ST
Russia
900 km, I = 90
3850
Baikonur
1.5 * 3.35
1
Vega
EU
700 km, I = 80
1580
ELA1
2.04 * 1.952
Soyuz-IkarFregat
Russia
700 km, I = 90
3000
Plesetsk
Proton-M
Russia
170 km, I = 72.7
19975
Baikonur
KVD - 1 (LH2LO2)
PS-4/L2
(MMH-Mon-3)
RL 10B-2
(LH2-LO2)
RL10B-2 (LH2LO2)
AJ 10-118 K
(AerozineN2O4)
RL-10A-3-3A
or RL-10A-4
(LH2-LO2)
27,2
30,69
Y
Y
20.410/27.200
23.260/3
0.690
Y
Y
6
6,95
Y
Y
16,74
18,8
Y
Y
RL-10A-4-2
(LH2-LO2)
20.672/20.830
22.586/2
2.936
Y
Y
YF-73 (LH2LO2)
YF-40 (UDMHN2O4)
S5-92 (UDMHN2O4)
8,5
10,5
Y
Y
14,15
15,15
Y
N
5,35
6,535
20
Y
0,418
(UDMH-N2O4)
0,55
0,968
5
Y
2.61 * 2.72 or
1.5 * 3.35
2.352/
1
UDMH-N2O4
0.3-0.9/5.35
3.29/6.53
5
20
Y
2.61 * 4.1
2,37
UDMH-N2O4
19,8
-
8
Y
Y
Y
Y
N
2
N
Falcon 1
US
700 km, I = 85
450
VAFB
-
H-2A
Japan
800 km, I = 98.6
4400
Tanegashim
10.7 * 4
Falcon 9
US
900 km, I = 80
7246
VAFB
-
CZ-3B/3C
China
800 km, I = 90
6000
XLSC
12.375 * 3
Soyuz-2
Russia
820 km, I = 98.7
4350/4900
Plesetsk
1.5 * 3.35
1
3
0,3
Proton-K
Russia
SSO
4600
Baikonur
5.5/6.3 * 3.7
2.500/
3.370
ROCKOT
Russia
800 km, I = 90
1340
Plesetsk
2.61 * 2.5
1,6
Kestrel (RP-1LO2)
LE-5B (LH2LO2)
Merlin (RP-1LO2)
YF-75 (LH2LO2)
S5-92 (UDMHN2O4)
Kerosene or
Sintin-LO2
11DM58
(UDMH-N2O4)
17
20
18,193
20,6
2
Y
5,35
6,535
20
Y
14,8
17.3/18.2
7
Y
4,9
6,5
8
Y
Soyuz Launch Vehicle
The Soyuz family of launch vehicles is the world’s most successful launch system, with more than
1200 launches and an overall success rate of about 96.8 %. It is based on the R-7 ICBM design and
has been in service since 1963. The current operational models are the Soyuz-U, Soyuz-FG, and
Soyuz 2. The Soyuz 2 vehicle is planned to replace all previous models by 2014. Hence, we will
focus on this vehicle.
The Soyuz 2 is a 4 stage launch vehicle, with an overall length of 43.42 m, maximum diameter of
10.3 m and a maximum lift off mass of 311.7 tons. It can launch up to 4.9 tonnes to a sunsynchronous orbit with an altitude of 820 km, inclination of 98.7 degrees from the Plesetsk
cosmodorme. Table 3 summarizes some of the characteristics of the soyuz 2 vehicle.
Table 3: Soyuz 2 characteristics
Stage
Length (m)
Max Diameter
(m)
Mass (kg)
Engine
Engine Mass
(kg)
Dry Mass (kg)
Propellant
O/F
Isp (sea level)
Isp (vacuum)
Propellant Mass
(kg)
Burn Time
Thrust (sea
level)
Thrust (vacuum)
Control
1
19.6
2
27.8
3
6.74
Fregat
1.5
2.68
2.95
2.66
3.35
176800
101900
RD-108A (+ 4
verniers)
25200
RD-0110/RD0124
6535
1155
1250
408/480
75
15240
2355
2.47
245
310
6875
Kerosene-LO2
2.39
264
311
2.2/2.6
331
326/359
1000
UDMH-N2O4
1.95-2.1
327 (225 verniers)
156640
90100
21380
5350
118
280-290
230-250 / 300
877 (total)
838.5 * 4
792.68
-
-
1021.3 * 4
990.18
245/294
19.6 + 7.8
(verniers)
Movable
aerodynamic fins
and 8 gimballed
verniers
(deviation angles
up to 45°)
4 gimballed
verniers
(deviation
angles up to 45°)
4 gimballed
verniers
(deviation angles
up to 40°)
-
4 * RD-107A
S5-92
An analysis of Soyuz launches during the last 3 years (2009-2011) was performed using online
resources which are summarized in Table 4. Not much data was available about non LEO launches.
For launches to the ISS, an elliptical parking orbit was used with an inclination of 51.62-51.68
degrees, a perigee of 189.6-201.2 km, an apogee of 232.41-267.02 km and a period of 88.47-88.87
minutes (7).
Table 4: Soyuz launch statistics 2009-2011 (8)
LEO/ISS
MEO
Molniya
LEO
Sun Synchronous
LEO
Number of
Launches
Launch Site
26
4
3
7
Baikonur
Plasetsks/Korou
Plasetsks
Baikonur/ Plasetsks
2
Baikonur/ Korou
Payload Mass
Minimum
Maximum
(kg)
(kg)
7.1
7.4
0.94
1.414
2
2
3.9
6.9
2.1
2.7
Figure 16: Fregat upperstage mission profile (SSO orbit from Guiana Space Center) (9)
Proton Family of Launch Vehicles
The proton family of launch vehicles is another successful Russian/Soviet launch system. It has
been in service since 1965, with more than 400 launches and a success rate of about 90%. The
Proton M vehicle is currently operational and we will focus only on this vehicle.
The Proton M vehicle is a 4 stage launch system, capable of delivering 19975 kg to a 170 km
circular orbit at an inclination of 75 degrees from Baikonur. It has a length of 57.2 m, a maximum
diameter of 7.4 m and a maximum lift off mass of 691.27 t. Table 5 summarizes the characteristics
of this vehicle. It should be noted that the Angara launch vehicle family which is going to replace
the proton vehicle in the future will be using the same upper stage.
Table 5: Characteristics of the Proton M launch vehicle (6)
Stage
Length (m)
Max Diameter
(m)
Mass (t)
Engine
Engine Mass
(kg)
Dry Mass (kg)
Propellant
O/F
Isp (sea level) (s)
Isp (vacuum) (s)
Propellant Mass
(kg)
Burn Time
Thrust (sea
level) (kN)
Thrust (vacuum)
(kN)
Control
1
21.18
2
17.05
3
4.11
Breeze M
2.61
7.4
4.1
4.1
4.1
30.6
RD-253
11.4
RD-0210
3.7
RD-0210
2.37
11 DM 58
1300
566
-
95
UDMH-N2O4
2.69
285
316
-
-
2
326.5
326.5
325.5
419410
156110
46560
19800
130
300
250
2320
583 (+31)
19.6
By gimballing four
Nozzles
By 4 verniers
engines (31 kN
thrust)
By 4 thrusters
(396 N thrust) and
12 attitude Control
thrusters
(13.3 N thrust)
10500
By gimballing six
nozzles
The proton launches for the last three years (2009-2011) were analyzed for patterns of target orbits.
The results are summarized in Table 6. The most detailed data existed for GEO launches by Proton
M. It revealed that typically a parking orbit with an altitude of 173 km, an inclination of 51.5
degrees and a RAAN of 9 degrees was used (10).
Table 6: Proton M launches (2009-2011) (8)
Number of
Launches
GTO
+
GEO
EEO
GTO
MEO
Launch
Site
14
6
1
5
5
Baikonur
Payload Mass
Minimum
Maximum
(kg)
(kg)
2.73
6.15
2.06
3.672
5.775
5.514
4.245
6.74
4.5
Figure 17: Typical Proton M mission profile (10)
Simulation of Launch
CHALLENGES


Detailed modeling of debris removal satellite (including propulsion subsystems) in STK:
Currently STK does not have this inbuilt capability. With the aid of the newly released
STK/SOLIS tool, it will be possible to address this challenge.
Modeling a tether system using STK with Astrogator and SEET module.
References
1. Instability of the present LEO satellite populations. Liou, J.C e Johnson, N.L. 7, 2008,
Advances in Space Research, Vol. 41, p. 1046-1053.
2. Chatters, E.P. e Crothers, B.J. Space Surveillance Network. AU-18 Space Primer. 2009.
3. United States Strategic Command: Space Track. [Online] http://www.space-track.org.
4. Collision activities in the future orbital debris environment. Liou, J.C. 9, 2006, Advances in
Space Research, Vol. 38, p. 2102-2106.
5. Satellite Collision Leaves Significant Debris Clouds. Orbital Debris - Quarterly News. s.l. :
NASA, 2009. Vol. 13.
6. European Space Agency. Launch Vehicle Catalogue. s.l. : European Space Agency, 2004.
7. History of Flights. S. P. Korolev Rocket and Space Corporation (Energia). [Online]
http://www.energia.ru/en/archive/launch-book.html.
8. Space Launch Log. [Online] http://www.spacelaunchreport.com/.
9. Soyuz from the Guiana Space Center User's Manual. 2006. 1.
10. Launch Information Processing and Display Center (LIPDC). Khrunichev (KhSC) State
Research and production Space Center. [Online] http://coopi.khrunichev.ru/main.php?id=11.
11. Isakowitz, Steven J., B., Hopkins Joshua e P., Hopkins Joseph. International Reference
Guide to Space Launch Systems. s.l. : American Institute of Aeronautics and Astronautics, 2004.
12. Characterization of the cataloged Fengyun-1C fragments and their long-term effect on the LEO
environment. Liou, J.C e Johnson, N.L. 2009, Advances in Space Research, Vol. 43, p. 14071415.
13. [Online] http://www.satellitedebris.net/whatsup/.