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Orbital Recovery's Responsive Commercial Space Tug for Life Extension Missions

  • Skycorp Incorporated

Abstract and Figures

Orbital Recovery Corporation (ORC) and its UK subsidiary Orbital Recovery Limited (ORL) are in the developmental stage of an orbital space tug called the Orbital Life Extension Vehicle (OLEV), whose purpose is to mechanically mate with an existing communications spacecraft in GEO or GEO intended orbit, take over north/south and east/west station keeping as well as attitude control. The OLEV is designed as a secondary payload on an Ariane V launch vehicle and carries a Hall Effect Thruster (HET) to execute GTO to GEO orbit raising, rendezvous and docking, and operations of the coupled spacecraft pair. The OLEV does not transfer fuel or otherwise interface with the parent spacecraft. The OLEV is designed to mate with any three axis stabilized spacecraft and has sufficient supplies to keep a 3000 kg parent spacecraft in geostationary orbit for up to an additional ten years of life.
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2nd Responsive Space Conference
April 19–22, 2004
Los Angeles, CA
Dennis Ray Wingo,
Orbital Recovery Corporation
Orbital Recovery Limited
London, UK
AIAA 2nd Responsive Space Conference
RS2 2004
AIAA 2ND Responsive Space Conference 2004
Copyright Orbital Recovery 2004, Published by AIAA 2nd Responsive Space Conference with Permission
AIAA-RS2 2004-3004
Dennis Ray Wingo, Orbital Recovery Corporation, London, UK
Orbital Recovery Corporation (ORC) and its
UK subsidiary Orbital Recovery Limited
(ORL) are in the developmental stage of an
orbital space tug called the Orbital Life
Extension Vehicle (OLEV), whose purpose is
to mechanically mate with an existing
communications spacecraft in GEO or GEO
intended orbit, take over north/south and
east/west station keeping as well as attitude
control. The OLEV is designed as a
secondary payload on an Ariane V launch
vehicle and carries a Hall Effect Thruster
(HET) to execute GTO to GEO orbit raising,
rendezvous and docking, and operations of the
coupled spacecraft pair. The OLEV does not
transfer fuel or otherwise interface with the
parent spacecraft. The OLEV is designed to
mate with any three axis stabilized spacecraft
and has sufficient supplies to keep a 3000 kg
parent spacecraft in geostationary orbit for up
to an additional ten years of life.
The life extension of GEO orbit spacecraft is
desirable to satellite operators for the
increased revenue potential that such life
extension provides. The proof principle of
this is the operation of numerous spacecraft
well beyond their contracted lifetimes. This is
typically accomplished by rationing the last
year(s) of station keeping fuel by eliminating
north/south station keeping which demands
ten times the impulse that east/west station
keeping requires (~40 meters/sec vs. ~3
meters/sec). This results in an inevitable
increase in orbital inclination due to the
gravitational influences of the Moon and Sun
at GEO altitude (~33,000 km). While the
increase in inclination requires spacecraft
users to purchase considerably more
expensive antennas to track the changing
inclination, the lower revenue charges by the
operators more than offset this cost. It is clear
that a great proportion of existing spacecraft in
GEO are actually in inclined orbits. Table 1
gives the numbers of GEO and inclined orbit
operational spacecraft as of June 2003.i
Region GEO Inclined
(160°W-71°E) 65 19
(71°E-0°E) 59 25
(0°W-61°W) 36 14
(61°W-160°W) 60 8
Table 1: GEO vs. Inclined Orbits
Out of a total of 286 operational spacecraft 66
of them are operating in inclined orbits. A
few of these spacecraft such as the old NASA
Tracking Data Relay Satellites (TDRS) are
used for communications with the South pole
installations during the descending nodes but
for the most part these inclined birds are used
for customers within the normal range of a
geostationary comsat. Some of these
spacecraft, such as the SBS IV/HGS-5,
launched on the Space Shuttle in 1984 (a
Hughes 376 spinner), are as much as nearly 20
years old and still functional and producing
revenues.ii The design life was 8 years.
The question that came to our minds was “can
we, in a cost effective manner, build a space
tug that could extend GEO assets lifetime and
keep them in GEO where their revenue
streams would be much higher than for the
AIAA 2ND Responsive Space Conference 2004
inclined birds?” Some companies make
considerable revenue from inclined birds but
their revenues are only a fraction of the
revenues obtained from operating in the
stationary belt.
In our business development we had to come
up with a metric to determine what cost
effective meant. Our design guidance was to
posit the simplest on orbit servicing system
possible and limit customers to those with the
most valuable commercial assets. The
spacecraft that fit these parameters are three
axis stabilized commercial GEO comsats with
a revenue of at least $40-50M dollars per year.
To us this meant that we had to achieve a cost
target for our system that was no more than a
year to a year and a half of revenue for the
operator. This would give an operator a
revenue of between $200-$500M over the life
extended mission between conservative and
optimistic assumptions. It turns out that out
of the existing 286 commercial comsats in
orbit today over 50 of them that fit our metrics
will need replacing by 2009.
We have to provide a clear value to the
customer and providing a three to six times
revenue return of our price qualifies. Business
with governments are handled on the same
basis as with commercial customers with
pricing based upon our commercial pricing
and not the other way around. Cost is the
fundamental driver of our business and if we
can not sell the OLEV for a commercially
attractive price that customers are willing to
pay then we do not have a business. Costing
like this for the commercial market is much
more understandable to investors as these are
metrics that anyone can analyze. Government
development efforts have different metrics
that are difficult for Wall Street to measure
and forecast. Government customers are an
additional source of revenue but our business
cannot be justified on that basis.
With cost efficiency as the principal system
architectural driver we undertook a detailed
design that worked for lower cost at each
stage. We decided that in order to meet any
reasonable overall cost target, we had to go
with a secondary launch to GTO. With
Arianespace we found a reasonably priced
secondary launch that went to our desired
orbit and with a payload weight of up to 1000
Kg. That put an upper limit on our size,
which ruled out chemical propulsion systems.
Time is also a factor and we needed a system
that was controllable in GEO in a robust
enough fashion to do the docking with the
only aperture available on the zenith face of a
commercial comsat, the liquid apogee motor.
In looking at available systems, one based
upon Hall Effect Thrusters (HET’s) seemed to
have better overall performance than gridded
ion in terms of handling the extra mass for
attitude control and station keeping of a
comsat as well as our asymmetrical load on
the parent spacecraft. This then drove the
size of the solar arrays to the 4-5 kilowatt
range in order to minimize transit time
between GTO and GEO.
In an unexpected development, during our
concept design study the Institute for Robotics
and Mechantronics DLR (German Space
Agency), demonstrated to us their advanced
development of a complete rendezvous and
docking solution for GEO, including a capture
tool that would allow us to dock to the apogee
motor and imaging software to guide the
system. We have an exclusive commercial
license for this solution and have integrated
this into our design. This agreement and the
hardware and software supplied allows us to
decrease schedule risk, increase our credibility
by leveraging the extensive design experience
at DLR, as well as allowing for a greater
definition of the cost and configuration of the
system early on in the design process.
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With these general conclusions we drew up a
strawman design and did preliminary costing
of subsystems and components to determine
raw costs. We added to this the expected cost
of manufacturing, launch, and insurance and
came up with a number within the cost cap
chosen as our market viability point. We then
rolled the requirements that came from our
study into a document that became the
foundation of our RFP to potential suppliers.
In the fall of 2003 after an evaluation period
for proposal responses, we chose Dutch Space
of Leiden, the Netherlands, because their
solution was almost ideal for our purposes.
The OLEV is based upon a modification of
the design of the cone adapter between the
2624 millimeter base of the Ariane V upper
stage and the standard 1194 millimeter
Marman adapter between the cone and the
primary payload(s). Figure one illustrates the
concept and placement within the Ariane V
Figure 1: OLEV on the Ariane V
The OLEV carries the standard Ariane V cone
adapter above its own internal structure. The
solar arrays and other deployables reside
underneath the spacecraft in an area that is a
spacer used for these secondary payload
missions. The solar array is a six-petal design,
carrying triple junction Emcore solar cells to
maximize power to approximately 4 kilowatts.
Adding more power is possible but this is
beyond our baseline model. Figure 2 shows
the available area inside the Ariane V adapter.
Figure 2: A 5 Secondary Payload Volume
This arrangement solves a host of problems
for Orbital Recovery’s application, including
providing a large surface area interface
between the OLEV and the parent spacecraft.
The spacecraft has a wide rear surface
necessary to provide a moment arm for the
station keeping thrusters, already incorporates
Hall thruster technology, (as a result of a prior
B0 study), and has the ability to incorporate
the Capture tool and rendezvous and docking
software from DLR. Figure three gives a
general picture of the layout of the spacecraft.
Figure 3: OLEV Internal Configuration
The OLEV carries the standard hardware
expected on a GEO comsat and is rated for a
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twelve year lifetime. This includes two
hundred and sixty kilos of xenon gas for the
GTO to GEO transfer and enough fuel to
station keep a 3000kg comsat mass for up to
ten years with margin. The attitude control
system is oversized in order to be able to
maintain the +/- 0.035 nadir pointing required
by the largest comsats. For electronics we are
using a spacecraft processor built by Swedish
Space, that already has proven much of our
orbit raising mission profile in the SMART-1
mission by ESA.
The Dutch Space OLEV design meets all of
our overall cost targets within an acceptable
margin. At the present time we have just
started a B1 design study in cooperation with
ESA under the ARTES4 public/private
partnership program whereby an ESA fund
matches the funding provided by the private
entities (Dutch Space and ORL). This fund is
set up to help to enable companies within the
European Union (EU) to generate new
aerospace business for EU based companies.
The robotic payload on the OLEV is being
supplied principally by DLR. This comprises
both hardware (capture tool) and software
(model matching software, sensor feedback
software from the capture tool, and
telepresence software) to allow ORL
engineers to guide the OLEV to a docking or
under autonomous operation. Figure 4 shows
the capture tool already built by DLR.
Figure 4: DLR Capture Tool
This capture tool will be flight qualified and
flight copies will be manufactured by Kayser
Threde, of Munich, Germany under a
commercialization agreement with the
German government. The flight and ground
software will also be supplied through Kayser
Threde from DLR. The capture tool houses
six sensor heads located at 120 degree angles
apart. These sensors give feedback to the tele-
operator or to the autonomous software to
determine where within the volume of the
apogee motor the capture tool is located. The
locking mechanism, when fully inserted into
the apogee motor crown, has a crown
mechanism that, using a set of spreading pins,
forms a tight mechanical connection between
the capture tool and the apogee motor. Figure
5 gives is an illustration of the capture tool
and sensors inside of an Apogee motor.
Figure 5: Capture Tool In Apogee Motor
The Dutch Space design of the OLEV
supplements this connection by pulling the
capture tool down to contact the ring or a plate
offset from the ring (avoiding the ejection
springs) in a manner under revision during the
B1 study.
The software from DLR acts as a feedback
mechanism between the sensors located in the
capture tool and the imaging system. The
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software works with a standard 2D imaging
system and uses a software model of the
parent spacecraft and pattern matches that to
the images received from the camera system.
Range and range/rate data can be generated by
the software based upon inputs from the
imager to allow the software to control the
rendezvous phase. Under consideration is a
laser rangefinder to supplement the imaging
system. Figure 6 is an example of the output
of the imaging and pattern matching software.
Figure 6: DLR Pattern Matching Software
DLR at their Institute for Robotics and
Mechatronics has a full simulation laboratory
where an extensive amount of work has been
carried out in developing the simulation for
this mission. The software acquired under
license by ORL is a major factor in reducing
our development time and will allow the team
led by Dutch Space to concentrate on
integrating the total system rather than
designing from scratch a very sophisticated
robotic hardware and software system.
The OLEV mission begins with the separation
of the spacecraft from the Ariane V launch
vehicle. Solar arrays and antennas are
deployed, initial system tests accomplished
and the Hall thruster system enabled. The
OLEV will take approximately 120-150 days
from separation to achieving GEO orbital
altitude. It will take approximately 120-140
kilos of Xenon to get to GEO. Figure 7 is a
representative transfer from GTO to GEO
simulated for us by SAIC and confirmed by
Dutch Space.
Figure 7: GTO to GEO Transfer
The transfer time is influenced by the beta
angle during the climb as well as the
orientation of the orbit with respect to the
Earth’s umbra and penumbra shadow. Also, it
is strongly desired to gain perigee altitude as
quickly as possible to get above 10,000 km.
Based upon data from the ESA Spacecraft
SMART-1 this is the altitude where solar
array degradation falls off to a low value.
The rendezvous and docking sequence begins
at an altitude slightly above GEO and behind
the parent spacecraft. This is the classical R
bar approach used by NASA in several
docking approaches with the Shuttle. The R
bar approach precludes plume impingement
on the parent spacecraft by the Xenon from
the Hall thrusters. Approach is accomplished
by using GPS or spacecraft orbital
determination and by knowing the location of
the parent spacecraft within a GEO box of 80
X 80 X 40 km. The spacecraft also knows the
attitude of the parent (usually nadir oriented to
within 0.7 degrees). This is the baseline that
may be refined during the B1 study period.
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After the OLEV closes to within 40 km visual
indications are used to guide the spacecraft
closer to the parent. It may be possible to use
the parent’s transponders as a homing beacon
or use a laser rangefinder at this distance.
After the parent is positively acquired and the
distance closes to less than 400 meters the on
board imaging system will be used to guide
the parent to docking. One dramatic
difference in performing a docking at GEO
altitude is that the order of magnitude of
forces (differential orbital velocity) is
approximately 256 times less than in a LEO
orbit such as the International Space Station.
The final stage is at 4 meters and closer where
the Hall effect thrusters may not have the
control authority to move the spacecraft
around. The Dutch Space led team in the B1
study is currently investigating a cold gas
system to control the spacecraft for the last 4
meters. The OLEV is oriented in a
nadir/zenith orientation and closure to within
the apogee motor is accomplished. The
LED’s indicate the depth within the nozzle
and when the crown mechanism has passed
the throat of the nozzle the crown locking
mechanism is actuated. After a positive lock
is accomplished the capture tool is retracted at
its base and latches are deployed from within
the OLEV to provide a three point positive
latch with the parent spacecraft Marman
clamp underside surface. Detailed design of
this mechanism is in progress. This provides a
multipoint interface that allows the capture
tool and the apogee nozzle to carry only a
fraction of the total loads of the coupled
system. Figure 8 gives an illustration of the
configuration of system just prior to docking.
Figure 8: OLEV Prior to Docking
At the moment of capture the on board
attitude and station keeping control system is
disabled. The OLEV then takes complete
control of both the attitude and station keeping
of the coupled spacecraft pair.
For the coupled system to correctly operate
the OLEV must dock to the center of mass in
the X/Y plane leaving only a Z offset to be
corrected. In order to provide maximum
revenue, the attitude control of the coupled
mass must be as good as for the original
system. For some spacecraft this is as fine as
+/- 0.035 degrees in all three axes. Station
keeping must also be accomplished for the
coupled mass for as long as ten years after
docking with enough reserve to allow for
moving to the final disposal orbit.
Station keeping for east/west station keeping
is accomplished by the use of the Hall
thrusters in the nadir or zenith direction,
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depending on the position in the GEO “box”
that the spacecraft wants to stay in. Station
keeping in the north/south direction has to be
through the center of mass of the coupled
system. The coupled system center of mass
shifts due to the depletion of the xenon fuel
during the life of the life extension mission.
This represents the most challenging technical
aspect of the OLEV design because an exact
representation of the parent spacecraft is not
known to sufficient precision prior to docking.
This will be compensated for by designing for
the worst case, with single or dual axis
gimbals on critical thrusters.
After a check out and training period the
operation of the coupled system is turned back
over to the satellite operator or their
designated contractor. Orbital Recovery does
not intend to operate customer satellites but
will provide a full compliment of engineering
backup should any problems occur. One
strategic consideration is that Orbital
Recovery will always retain the capability to
undock the spacecraft from the parent comsat.
This is in the case of missions where the
comsat dies before the exhaustion of xenon
fuel in the OLEV and we have delivered the
defunct spacecraft to the junk orbit. The
customer shall not have the ability to do this,
only Orbital Recovery, in order to absolutely
maintain control of the operation of the system
when in free flight mode.
The OLEV has great potential as a responsive
space system for commercial as well as
government customers. In the commercial
realm, in space insurance is at a historic high
today. We are in discussions with a potential
customer to provide a OLEV to dock with a
retiring commercial asset that has been
replaced with a new model of similar
specifications. This gives the customer
effectively a “one spacecraft deductible” on
their yearly insurance. In effect if the primary
spacecraft fails for any reason, the customer
simply brings the retired spacecraft back
online. Basically this is a hot spare that would
cause an outage of a few hours versus the long
periods of time and the expensive acquisition
of spare capacity on a competitor’s spacecraft
Another method for providing responsive
space capacity for commercial fleet operators
is to provide a OLEV on orbit to backup any
fleet operator’s propulsion related difficulty.
This capability results in lower insurance rates
and a higher level of customer availability.
This same rational could easily apply to
government customers. At this time there are
several semi-retired TDRS spacecraft in
inclined orbits that could be rehabilitated or as
existing operational assets are retired they
could be life extended as either a hot spare or
as an augmentation of the existing fleet.
NASA at the current time is having a great
deal of difficulty related to ISS
communications capability. A system such as
ours would solve this problem at a fraction of
the cost of augmenting the existing TDRS
For other government customers in GEO
similar benefits would accrue. The USAF is
in the midst of a massive upgrade in
communications capability via the
Transformational Communications
Architecture. Recent reports indicate that this
program, as well as the Advanced EHF system
are having schedule and cost issues. A OLEV
could dock with an existing Milstar spacecraft
and extend its life to compensate for shortfalls
in expected AEHF production, or to keep the
multi billion dollar Milstar spacecraft in
service beyond their expected fuel lifetime.
The cost metrics that make our system a
favorable alternative in the commercial arena
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are multiplied by the ratio of the cost
difference between the OLEV and the
spacecraft that they extend the life of. Figure 9
is an illustration of the OLEV as a free flyer
being deployed from an Ariane V.
Figure 9: OLEV Post Deployment
Going a step beyond simple life extension
commercial customers in the future as well as
the U.S. government could buy OLEV
vehicles and have them on station above the
GEO belt or below the GEO belt to rapidly
respond to failures of upper stages or the on
orbit propulsion systems of their assets. The
government in operational structure has the
same issues as a commercial fleet operator.
The OLEV could be placed into orbit on
contingency in order to provide rapid response
capability to recover a critical asset who’s life
has been curtailed by a propulsion related
Today in GEO orbit there are no less than
seven commercial spacecraft with recent
partial or total failures of the propulsion
system that could be mitigated by the use of
the OLEV. Two years ago a delivery of
TDRS-I to GEO orbit from GTO was delayed
by several months due to a flaw in the
spacecraft propulsion system. Before that, the
European Artemis spacecraft was left in an
improper transfer orbit between GTO and
GEO by a failure of the Ariane V upper stage.
On that same flight an Orbital Sciences light
GEO sat was left in an unusable orbit without
the propulsion capability to achieve GEO with
any usable life. Since then the 5000 kilogram
Astra 1K was stranded in a LEO orbit by the
failure of the upper stage of the Proton rocket.
In the late 1990’s a U.S. Navy spacecraft was
left in an unusable orbit. On average an upper
stage propulsion failure occurs every 18
months. The OLEV could also be placed on
station in GEO to respond within a couple of
weeks to a problem that could cripple a
critical space asset during a period of
increased operational tempo. Also, the
existence of contingency OLEV spacecraft
would provide considerable increased
operational flexibility in mission planning by
allowing planning beyond the depletion of the
existing on board fuel supply of multi-billion
dollar communications assets.
The OLEV could also be used for responsive
operation with large LEO government assets.
In general, the OLEV is too expensive for
commercial LEO assets but it would be
suitable and cost effective to add life or
contingency propulsion for expensive
government LEO assets. Beyond GEO, a
OLEV could be used by NASA in the case of
the failure to deploy of the petals of the James
Webb space telescope.
For other applications the OLEV has the
potential to grow into a very capable
responsive in-space servicing system.
Discussions are underway now with potential
customers for providing power for the Boeing
702 spacecraft with large solar array
degradation problems. As the OLEV matures
it may be possible to change the way that
GEO spacecraft are procured in order to
facilitate their servicing by the OLEV Mark II
system that could replace components,
provide extra power, or life extension. Going
beyond this even it can be forecast that the
existence of proven, cost effective on orbit
servicing spacecraft could change the market
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metrics for GEO spacecraft by allowing the
designs to be less long lived, with servicing
built in. This would reduce cost dramatically
for the manufacturer and the customer. The
overall point here is that the mere existence of
a proven Orbital Recovery product in this
market will begin to open the door to new
ideas and applications that formerly have been
beyond the grasp of the existing way of doing
business. The key is cost effectiveness and
cost effectiveness is enforced by the
commercial market that has little stomach for
expensive solutions and long term R&D
projects. Making money is the key parameter
for success in commercial space. With our
success we can help make space more
profitable for commercial interests and more
responsive for everyone.
... A binocular vision strategy based on image matching is developed for capturing noncooperative spacecraft [1]. A binocular extra-close relative pose measurement method based on geometry matching is proposed [17]. In [14,15], a position and pose measurement method based on box model matching is proposed. ...
... After applying gradient calculation, the edge extracted from the gradient value is still blurred. With respect to criterion 3 [17], there should only one accurate response to the edge. Thus, non-maximum suppression is used to suppress all the gradient values (by setting them to 0) except the local maxima, which represents the points with the sharpest change of intensity value. ...
... The value of A, B, C can be solved by the least squares method [17] and we obtain vector OZ * as: ...
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For finding an effective measurement for space non-cooperative targets in close distance, a vision-based position and pose determination algorithm of non-cooperative target for on-orbit servicing is investigated. First, the satellite-rocket docking ring’s region is separated from the background environment by a selected grayscale thresholding method and the Canny operator algorithm is used to extract the edge of the satellite-rocket docking ring. Furthermore, the mathematical alignment of the image plane and binocular stereo matching are used to obtain the correspondence of 3D points. Finally, the model of 3D reconstruction is established to determine the relative position and pose of the satellite-rocket docking ring. The ground simulation test shows that our method can measure the relative position and pose of the satellite-rocket docking ring effectively. The precision of the relative position is 0.001 m and that of the relative attitude is 0.1 deg., which satisfies the requirements of the relative measurement for non-cooperative target in extra-close range.
... The space robots (space remote manipulator system), such as B Weizhong Guo 1 School of Mechanical Engineering, Shanghai Jiao Tong University, Shanghai, China Canadarm, Dextre, and ETS-VII [1], have been a crucial element in all space construction activities and other space missions. The space missions currently being planned by space agencies around the world, like Orbital Express mission of the U.S. Defence Advanced Research Project Agency (DARPA) [2] and the ConeXpress Orbital Life Extension Vehicle (CXOLEV) of Orbital Recovery [3], show an increase in the number of robots. ...
... We will discuss this problem in our later works. The origin of moving coordinate system C M with respect to base coordinate system C B is represented by the position vector r = [x O y o z o ] T , while the orientation of moving coordinate system C M with respect to base coordinate system C B is represented by the rotation matrix R ∈ SO (3). The homogeneous transformation matrix describing the pose of the manipulator is given by ...
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A new remote manipulator based on cable-driven parallel mechanism (CDPM) is designed for space long-distance operations (e.g. space capture/docking and other long-distance space activities) in this paper. By controlling the cables and thrusters which are equipped on the manipulator simultaneously, the new remote manipulator can achieve expected position, linear velocity, and angular velocity. The new manipulator has a larger controllable workspace compared with usual CDPMs. The structure and characteristics of this manipulator are discussed in this paper. The volume and characteristics of the workspace are also discussed. The influence of the distance on the static equilibrium is studied. The simulation results show that the workspace of this new manipulator is larger than usual CDPM’s. The results also indicate that the cable forces and thruster vectors can completely constrain the manipulator and meet the requirements of space activities. The results of the simulation also show that the controllable workspace of the manipulator is not continuous at some regions. Hence, trajectory planning is necessary.
... It comprised a spacecraft based on the SMART-1 bus to mechanically latch onto a client satellite at the apogee engine nozzle to provide orbit transfer and/or station-keeping functions using electric propulsion-it required no interfacing with the client spacecraft which operates normally other than the mechanical dock with its apogee engine capture tool. It had heritage from the earlier cancelled ConeXpress [16] (formerly SLES [17]), a service module to provide 12 years of extended life support to aging geostationary satellites. Using Hall thrusters, it would dock with the target satellite until the capture tool on a retractable boom was inserted into the thrust cone of the inert apogee kick engine without the use of robotic manipulators. ...
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Space-based manipulators have traditionally been tasked with robotic on-orbit servicing or assembly functions, but active debris removal has become a more urgent application. We present a much-needed tutorial review of many of the robotics aspects of active debris removal informed by activities in on-orbit servicing. We begin with a cursory review of on-orbit servicing manipulators followed by a short review on the space debris problem. Following brief consideration of the time delay problems in teleoperation, the meat of the paper explores the field of space robotics regarding the kinematics, dynamics and control of manipulators mounted onto spacecraft. The core of the issue concerns the spacecraft mounting which reacts in response to the motion of the manipulator. We favour the implementation of spacecraft attitude stabilisation to ease some of the computational issues that will become critical as increasing level of autonomy are implemented. We review issues concerned with physical manipulation and the problem of multiple arm operations. We conclude that space robotics is well-developed and sufficiently mature to tackling tasks such as active debris removal.
... Recent studies concern the use of unmanned spacecraft in orbital servicing missions, such as the on-orbit refueling of telecommunication satellites, to extend their operational life (the operational life of many satellites is usually limited due to expired propellant, required for station-keeping) or space tugs [8][9][10]. In both cases, since such spacecraft are unmanned, a robust flight software must be developed. ...
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This paper outlines a method based on the theory of artificial potential fields combined with sliding mode techniques for spacecraft maneuvers in the presence of obstacles. Guidance and control algorithms are validated with a six degree-of-freed (dof) omorbital simulator. The idea of this paper is to provide computationally efficient algorithms for real time applications, in which the combination of Artificial potential field (APF) and sliding mode control shows the ability of plan trajectories, even in the presence of external disturbances and model uncertainties. A reduced frequency of the proposed controllers and a pulse width modulation (PWM) of the thrusters are considered to verify the performance of the system. The computational performance of APF as a guidance algorithm is discussed and the algorithms are verified by simulations of a complete rendezvous maneuver. The proposed algorithm appears suitable for the autonomous, real-time control of complex maneuvers with a minimum on-board computational effort.
... Rendezvous and Docking in GEO is envisaged for the servicing of communication satellites, e.g. [de Peuter et al. 1994], [Wingo 2004], [Creamer et al. 2006], [Kaiser et al. 2008]. It is a novel operation, not yet done, and there is little information available concerning problems, such as the effect of orbital disturbances on the rendezvous trajectories. ...
Conference Paper
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Whereas a lot of experience has been gained with rendezvous and docking operations in Low Earth Orbit (LEO), in Geostationary Orbit (GEO) no such operation has been performed so far. One of the differences to rendezvous in LEO is in GEO the disturbance of the rendezvous trajectories by the solar pressure, the direction of which changes by 360 deg during an orbit (1 day). The paper analyses for various starting times during the day the evolution of the position of a spacecraft w.r.t. an undisturbed local coordinate frame as a result of the solar pressure. It further analyses how two spacecraft with different ballistic coefficients are moving relative to each other under the influence of solar pressure. Impulsive radial-and tangential boost trajectories and constant z-and x-force trajectories are investigated for a typical transfer range in the close range rendezvous phase of the order of a few hundreds of metres. Conclusions concerning the suitability of various trajectory types for the close range rendezvous and concerning trajectory safety in GEO are drawn.
The concept of the OOS of spacecraft can be traced back to the 1960s, when the main focus was on providing the necessary maintenance to advance the lifetime of spacecraft and extending the scale and function through on-orbit assembly. During the past decades, the Hubble Space Telescope has made great contributions to the fields of astronomy and physics through both observational data and the success of five OOS missions to overcome big challenges. That included an initial flaw of its primary mirror and subsequent obstacles associated with replacing and upgrading its science instruments. Furthermore, many programs have been carried out in the area of the OOS of spacecraft with successful operations in space. It could be exemplified by the assembly of the International Space Station, service verification of ETS-VII and Orbital Express, and detailed research for future applications including servicer and client satellites and particularly large space systems. This paper attempts to summarize all reported programs of the OOS in terms of engineering developments and provide an overall perspective for investigators in this field. Based on the reviewed programs, an analysis is carried out to elucidate the logical architectures of the mission and technology of the OOS of spacecraft. Further attention is paid to discussions of the enabling technologies that support the development of the OOS and related spacecraft. As an outlook, the future development and challenges of the OOS and the application of novel technologies are finally discussed to extend the present review work.
Conference Paper
This paper investigates the synchronized position and attitude tracking control of a spacecraft tracking a space object. The objective of the control is to make the spacecraft stay at a certain distance above a specific surface of the space object. Dual quaternion is utilized to describe the spacecraft’s orbit and attitude coupled equations of motion. A generalized PD controller is developed to fulfill the control objective. Stability is proved by the Lyapunov’s stability theory and the Barbarat’s lemma. Numerical simulations illustrate the effectiveness of the tracking control.
When a service spacecraft docks successfully with its target spacecraft and forms a docked spacecraft, it will cause a large shift in the dynamics of the docked spacecraft. Not only do the mass properties change, but so do the reaction wheels' configuration. Meanwhile, the attitude of the docked spacecraft will inevitably change, under the influence of contact and impact, and it may lead to instability of the entire system. Due to the limited control torque and saturation, the control system of the reaction wheel may not guarantee the system stability, and the thruster can generate large control torque, but it consumes valuable jet fuel. Since the space manipulator may generate a greater coupling torque by movement, this paper proposes an attitude coordinated control method for docked spacecraft based on the estimated coupling torque. The method adopts the chaotic particle swarm optimization (CPSO) algorithm to plan the coordinated motion trajectory of space manipulator, and then designs a coordinated control law based on the estimated coupling torque of space manipulator to achieve the attitude control of docked spacecraft in order to guarantee the system stability. Numerical simulations validate the feasibility of the proposed method. In comparison with the traditional attitude control method, the attitude coordinated control method makes use of the coupling torque of the space manipulator, and overcomes the shortcomings of the limited control torque and saturation of reaction wheel, without consuming expensive jet fuel.
Solar power has merit as a renewable source of energy; it is the largest asset available for consumption on Earth and is limitless. There have been many ideas proposed to beam solar power to Earth; all have been dependent upon the provision of a backing frame to support solar panels, photovoltaic cells, and transmission. This paper suggests one type of rigid deployable skeletal structure and its material of manufacture to form the backing frame of solar panel systems; the structure takes the form of a skeletal double-layer tetrahedral system. The composite material is polyethersulphone thermoplastic polymer reinforced with carbon fiber (CF). This paper also discusses the hostile environment of space in relation to the fiber-reinforced polymer (FRP) composite material, with special reference to the thermal response. Finally, it suggests the alternate possibility of using a rigidized inflatable flexible skeletal structure and, as far as is possible, compares (a) the relative cost of transferring the two structures to low Earth orbit from Earth, and (b) the cost of solar energy relative to other forms of energy.
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