Note: Descriptions are shown in the official language in which they were submitted.
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SYSTEM AND METHOD FOR UNLOADING ANGULAR MOMENTUM FROM A
SPACECRAFT MOMENTUM WHEEL STABILIZATION SYSTEM
FIELD OF THE INVENTION
The present invention relates generally to the positional control of
spacecrafts. More
particularly, the present invention relates to angular momentum unloading from
a spacecraft.
BACKGROUND OF THE INVENTION
Because of gravitational effects from the earth, sun and moon and, because of
the
solar photon flux, geosynchronous spacecraft require periodic maneuvering to
maintain the
spacecraft properly located and oriented with respect to the earth. Maneuvers
regarding the
spacecraft location are known as station-keeping maneuvers while maneuvers
regarding the
orientation or attitude of the spacecraft can include momentum unloading
maneuvers. Both
types of maneuvers require judicious use of spacecraft thrusters.
Keeping the spacecraft in a pre-determined orientation is made possible by
having a
momentum wheel stabilization system on board the spacecraft. Such
stabilization systems are
known in the art and generally comprise spinning flywheels defining roll,
pitch and yaw
angular momentum axes and a closed loop control system for controlling the
angular velocity
of the flywheels in accordance with external forces applied to the spacecraft.
The flywheels
are limited in their angular velocity and consequently require periodic
angular momentum
unloading if the spacecraft wants to maintain a fixed orientation with respect
to the earth.
Different approaches to station-keeping and angular momentum unloading
maneuvers are
known and involve the firing of the spacecraft thrusters following a detailed
calculation of the
thrusters required orientation and burn time, which can vary substantially
depending on the
type of thruster carried by the spacecraft. For example, a station-keeping
maneuver involving
bi-propellant thrusters can last 5 minutes, while a similar maneuver using ion
propulsion
thrusters, such as xenon ion propulsion system (XIPS), can last 3 or 4 hours.
Station keeping and momentum unloading maneuvers can be planned in advance and
performed in accordance with a schedule. Additionally, sensing and control
systems on
current spacecrafts are advanced enough to adapt the scheduled maneuvers to
changes in
accordance with external forces acting on the spacecraft. However, important
changes in
external forces can occur between scheduled maneuvers and require a response
before the
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next scheduled maneuver. This is particularly true for angular momentum
unloading
maneuvers.
With respect to angular momentum unloading operations using XIPS thrusters,
the
manufacturers of such thrusters will usually provide their customers with
proprietary
software for calculating the orientation and bum time of thrusters required to
unload desired
angular momentum. Typically, in a spacecraft equipped with four XIPS thrusters
disposed on
a same side of the spacecraft and each attached to a gimbaled arm system
having two degrees
of freedom, the manufacturer supplied program will search for a pair of
diagonally opposed
thrusters that can be suitably oriented and fired sequentially to effect the
desired angular
momentum unloading while minimizing the amount of drift imparted to the
spacecraft.
Additional information regarding on-orbit station keeping of satellites
equipped with XIPS
thrusters can be found in the papers: "On-Orbit Stationkeeping With Ion
Thrusters: Telesat
Canada's BSS-702 Experience" by T. Douglas et al., Proceedings of the Eighth
International
Conference On Space Operations, Montreal, Canada, May 17-21, 2004, and
"Performance of
XIPS Electric Propulsion in On-Orbit Station Keeping of the Boeing 702
Spacecraft" by
D.M. Goebel et al., Proceedings of the 38" IAA/ASME/SAE/ASEE Joint Propulsion
Conference, 7-10 July 2002, Indianapolis, Indiana, U.S.A.
Although such two-thruster approaches are in most cases adequate, they can
last many
hours thereby consuming significant electrical energy, since XIPS thrusters
must produce
2o electric fields to function. This can be disadvantageous when electrical
energy is in limited
supply, such as when the spacecraft is in the earth's shadow. Additionally,
such
manufacturer-supplied programs tend to offer little or no practical guidance
to the satellite
operator of what maneuvers are being effected. The supplied program provides
little more
than a series of operational parameter values to be uploaded to the
spacecraft.
Therefore, it is desirable to provide an improved system and method for
unloading
angular momentum from a spacecraft.
SUMMARY OF THE INVENTION
It is an object of the present invention to obviate or mitigate at least one
disadvantage
of previous approaches to unloading angular momentum from a spacecraft.
In a first aspect of the invention, a method is provided for unloading angular
momentum from a spacecraft based on desired angular momentum corrections, each
desired
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momentum correction being associated with a spacecraft momentum axis. The
method
comprises selecting a thruster and a desired thruster burn time. For each
desired momentum
correction, generating a constant angular momentum mapping of values of a
first orientation
parameter of the thruster to values of a second orientation parameter of the
thruster. A
specific orientation of the thruster is then selected in accordance with the
mappings.
In a second aspect of the invention, a system for determining an orientation
of a
thruster used in unloading angular momentum from a spacecraft is provided. The
system
comprises a calculation module, and an analysis module. The calculation module
is operable
to generate, in accordance with at least one pre-determined thruster burn time
and desired
momentum corrections, a constant angular momentum mapping of values of a first
orientation parameter of the thruster to values of a second orientation
parameter of the
thruster for each desired momentum correction and for each pre-determined
thruster burn
time. The analysis module is operable to analyze the mappings.
In a third aspect of the invention, a method for selecting a single thruster
for an
angular momentum unloading operation from a spacecraft is provided. The method
comprises: inputting desired angular momentum corrections to a calculation
module;
generating, for each of at least two thrusters, a constant angular momentum
mapping of
values of a first orientation parameter for each of the at least two thrusters
to values of a
second orientation parameter for each of the at least two thrusters for each
desired angular
momentum correction for a pre-determined thruster bum time; and selecting one
of the at
least two thrusters in accordance with the constant angular momentum mappings
of the at
least two thrusters.
In a fourth aspect of the invention, a system for selecting a thruster for an
angular
momentum unloading operation from a spacecraft is provided. The system
comprises a
calculation module and analysis module. The calculation module is operable to
generate, in
accordance with at least one pre-determined thruster burn time, and desired
momentum
corrections, a constant angular momentum mapping of values of a first
orientation parameter
to values of a second orientation parameter for each desired momentum
correction, for each
pre-determined thruster burn time and for each of at least two thrusters. The
analysis module
is operable to analyze the mappings.
In a fifth aspect of the invention, a system for unloading angular momentum
from a
spacecraft is provided. The system comprises: a gimbaled thruster, a thruster
controller and a
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sub-system having a calculation module and an analysis module. The gimbaled
thruster is
attached to the spacecraft and positionable relative thereto. The sub-system
is operable to
determine an orientation of the gimbaled thruster. The calculation module is
operable to
generate, in accordance with at least one pre-determined thruster burn time,
and desired
momentum corrections, a constant angular momentum mapping of values of a first
orientation parameter of the gimbaled thruster to values of a second
orientation parameter of
the gimbaled thruster for each desired momentum correction and for each pre-
determined
thruster burn time. The analysis module is operable to analyze the mappings to
determine an
orientation of the gimbaled thruster to unload desired angular momentum from
the spacecraft;
and the thruster controller is operable to receive the orientation of the
gimbaled thruster from
the sub-system.
In a sixth aspect of the invention, a system for visualizing orientation
parameters of a
thruster used in an angular momentum unloading operation of a spacecraft is
provided. The
system comprises a calculation module, a graph module, an analysis module and
a display
module. The calculation module is operable to generate, in accordance with at
least one pre-
deterniined thruster bum time, and desired momentum corrections, a constant
angular
momentum mapping of values of a first orientation parameter of the thruster to
values of a
second orientation parameter of the thruster for each desired momentum
correction and for
each pre-determined thruster burn time. The analysis module is operable to
analyze the
mappings. The graph module is operable to generate a graph of the first
thruster parameter as
a function of the second thruster parameter for each mapping; and the display
module is
operable to display the graph.
In a seventh aspect of the invention, an apparatus for unloading angular
momentum
from a spacecraft based on desired angular momentum corrections, where each
desired
momentum correction is associated with a spacecraft momentum axis, is
provided. The
apparatus comprises means for selecting a thruster and a desired thruster burn
time; means for
generating a constant angular momentum mapping of values of a first
orientation parameter
of the thruster to values of a second orientation parameter of the thruster
for each desired
angular momentum correction; and means for selecting a specific orientation of
the thruster
in accordance with the mappings.
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Other aspects and features of the present invention will become apparent to
those
ordinarily skilled in the art upon review of the following description of
specific embodiments
of the invention in conjunction with the accompanying figures.
BRIEF DESCRIPTION OF THE DRAWINGS
Embodiments of the present invention will now be described, by way of example
only, with reference to the attached figures, wherein:
Fig. 1 shows a plan view of a satellite having four thrusters;
Fig. 2 shows a satellite angular momentum stabilization system and controller
together with a thruster controller;
Fig. 3 shows a communication system of satellite control centre;
Fig. 4 shows an embodiment of the system according to an embodiment the
present
invention;
Fig. 5a is a graphical representation of a mapping of torque with respect to
thruster
orientation;
Fig. 5b is a graphical representation of constant angular momentum mappings;
Fig. 6 is a graphical representation of three intersecting constant angular
momentum
mappings;
Fig. 7 shows a graphical user interface according to an embodiment of the
present
invention;
Fig. 8 is a depiction of a centroid of a closed area surface defined by three
intersecting
constant angular momentum mappings;
Fig. 9 is a depiction of a point of minimization of an error related to a
closed area
surface defined by three intersecting constant angular momentum mappings; and
Fig. 10 is a flow chart of a method of an embodiment of the present invention.
DETAILED DESCRIPTION
For the purpose of the description, the terms satellite and spacecraft are
interchangeable. Generally, the embodiments of the present invention provide a
method and
system for performing an angular momentum unloading operation of a satellite
equipped with
dynamically oriented thrusters. According to embodiments of the invention, the
angular
momentum unloading operation can be performed with a single thruster, as
opposed to two
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diametrically opposed thrusters, as is common in previously known systems and
methods.
The single thruster approach can be an improvement over two-thruster
approaches in that it
provides the user with the flexibility to choose a desired thruster burn time,
thereby reducing
the amount of energy required in the angular momentum unloading operation.
This is can be
advantageous when the thrusters used in the unloading operation are thrusters
operating on
electricity, such as, for example, XIPS thrusters. According to embodiments of
the invention,
the system and method can be less energy intensive and can provide the
satellite operator
control over the power usage during angular momentum unloading operations.
Additionally,
embodiments of the system can provide the satellite operator with an analysis
module and
interface, which can operate in real-time or in near real-time, allowing the
operator to
consider alternate scenarios for unloading angular momentum from the
spacecraft.
Fig. 1 is a plan view of a satellite 20 having four thrusters 22 attached to a
bus 24.
Although four thrusters are depicted, the present invention can apply to a
satellite having any
suitable number of thrusters. Solar energy conversion panels 26 attached to
the satellite 20
are also shown.
With reference to Fig. 2, satellites, such as satellite 20, are generally
equipped with an
angular momentum wheel stabilization system 28 connected to a momentum
controller 32.
The momentum controller is connected to sensors 30 that can detect changes in
the attitude of
the satellite 20, the changes in the attitude attributable to gravitational
forces and/or to the
solar photon pressure. In some non-limiting embodiments of the present
invention, the
sensors 30 can detect even minute changes in the attitude of the satellite 20.
Upon such
attitude changes being detected, the momentum controller 32 passes values of
the attitude
variation to a processor 34, which calculates the changes required in the
angular momentum
stabilization system 28 to counterbalance the attitude changes in the
satellite 20. In practice,
the changes in the angular momentum stabilization system 28 include changes in
the angular
velocity of spinning flywheels (not shown) comprised in the angular momentum
stabilization
system 28. Angular momentum stabilization systems can include three flywheels,
each
flywheel defining an angular momentum axis. Thus, the angular momentum
stabilization
system and the satellite 20 have associated roll, pitch and yaw angular
momentum axes.
Once the required changes in the angular velocity of the flywheels have been
calculated, the momentum controller adjusts the angular velocities of the
flywheels
accordingly and the sensors 30 pass on updated values of attitude changes of
the satellite 20.
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1 .
Thus, the sensors 30, momentum controller 32 and the angular momentum
stabilization
system 28 define a feedback loop.
In the case where repeated corrections in the angular velocities of the
flywheels are
effected to correct for repeated external forces acting on the satellite 20 in
the same direction,
the angular velocity of the flywheels will keep increasing. Angular momentum
stabilization
systems are generally limited in the maximum allowable flywheel angular
velocities. Thus,
without effecting an angular momentum unloading operation, the attitude of the
satellite 20
can become compromised. As is known in the art, angular momentum unloading
operations
involve the firing of thrusters 22 in order to generate a force on the
satellite that will cause the
flywheels to decrease their angular velocity. The parameters of an angular
momentum
unloading operation are usually determined at a satellite control centre.
In satellites in general, the momentum controller 32 is connected to a
communication
interface 36, which is connected to a transceiver unit 40 in communication
with the satellite
command centre. The momentum controller 32 passes the values of the roll,
pitch and yaw
angular momentums to the satellite command centre via the communication
interface 36 and
the transceiver 40. Fig. 3 shows a satellite control center 41 in
communication with the
satellite 20. The measured roll, pitch and yaw momentums of the angular
momentum
stabilization system 28 of the satellite 20 are transmitted to the satellite
control center 41.
Based on these angular momentum values, parameters of an angular momentum
unloading
operation are determined. In an embodiment of the invention, the parameters of
the angular
momentum unloading operation are determined as follows.
Telemetry data of the measured roll, pitch and yaw angular momentum values of
the
satellite 20 are received at the satellite control centre 41 where a satellite
operation protocol is
used to analyze these received angular momentum values and determine if an
angular
momentum unloading operation is required. The satellite operation protocol can
involve a
satellite operator analyzing the received values and, based on experience,
deciding whether or
not corrective measures are required. Alternatively, the satellite operation
protocol can
include an automated computerized analysis as would be understood by a worker
skilled in
the art. If a corrective measure is warranted, desired angular momentum
corrections are
determined. In the case where the total angular momentum is to be brought down
to zero, the
desired angular momentum corrections are simply the negative of the measured
angular
momentum values. In any event, once the desired angular momentum corrections
have been
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determined, they are input into input module 48 of the computation system
depicted in Fig. 4.
Further, if the angular momentum behavior between angular momentum unloading
operations is predictable, the satellite operator can choose to bias the
angular momentum of
the angular momentum stabilization system in such a way that the angular
momentum
components values will go through zero between unloading operations. This can
be
accomplished, for example, by setting a desired angular momentum correction
value to a
measured angular momentum value plus a bias factor, where the bias factor is
the same parity
as the measured angular momentum value. As will be understood by a worker
skilled in the
art, the computation system can be part of the satellite control center 41 as
shown in Fig. 4.
lo Alternatively, the computation system can be separate from the satellite
control centre 41 but
connected to it through a network infrastructure (not shown).
Detailed characteristics of the satellite 20 are programmed into a calculation
module
50. Such characteristics include the satellite's dimensions, weight,
distribution of weight,
center of mass, thruster location, thruster orientation parameters, fuel
level, etc. These
characteristics are used to calculate the parameters of the angular momentum
unloading
operation. The desired angular momentum corrections entered at the input
module 48 are
passed to the calculation module 50. Additionally, an identification (ID) of a
single thruster
22 to be used in the angular momentum unloading operation and a thruster burn
time can be
entered into the calculation module 50 via an input module 52. Alternatively,
these thruster
ID and thruster burn time can be pre-programmed into the calculation module
50.
The present method and system result from a recognition that, contrary to
previous
belief, a single thruster can be used to effect an angular momentum unloading
operation,
provided the resulting unloading operation will not impart significant drift
to the satellite. As
such, thrusters having low thrust, such as XIPS thrusters, are ideally suited
to such single
thruster operations. High thrust thrusters, such as bi-propellant thrusters,
although they might
be used, are not ideal for such single thruster maneuvers since their thrust
is too great and
may result in substantial drift of the satellite. The fact that current
angular momentum
unloading operations are carried out using more than one thruster appears to
be historical. At
the onset of man-made satellites in space, only powerful thrusters were
available and angular
momentum unloading operations led to the use of two or more thrusters in such
corrective
procedures in order to unload angular momentum without imparting substantial
drift to the
satellite. While thruster technology has evolved, and lower thrust thrusters
are now available,
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their use for unloading angular momentum has not kept pace with the
technological changes,
resulting in energy-inefficient operations.
Upon receiving the required inputs, the calculation module 50 can effect a
mapping of
the selected thruster 22 torque about each of the roll, pitch and yaw axis of
the satellite 20 as
a function of a first thruster orientation parameter and of a second thruster
orientation
parameter. As an example, in the case where the thruster 22 can be oriented by
the selection
of two angles, Fig. 5a shows a graphical representation of such a mapping of
torque values
on surface 61, for torque about the roll axis as a function of a first angle
(p) and a second
angle (y).
As will be understood by a worker skilled in the art, selecting a torque value
and
multiplying it by the burn time of the thruster 22 yields the angular momentum
effected by
the thruster 22. Thus, upon mapping the roll, pitch and yaw torques as a
function of the
orientation angles of the thruster 22, the calculation module 50, based on the
selected thruster
22 burn time, can effect a mapping of values of a first thruster orientation
parameter to values
of a second thruster orientation parameter for each of the roll, pitch and yaw
desired angular
momentum corrections. Fig. 5b shows, as an example, a graphical representation
of such a
mapping for a desired roll angular momentum correction of -10 Nms and a pre-
determined
bum time of 1500s. Fig. 6 further shows, as examples, additional pitch and yaw
graphical
representation of mappings for a desired pitch angular momentum correction of
6 Nms and a
desired yaw angular momentum correction of 3 Nms for the same pre-determined
burn time
and specific thruster.
In general cases, as shown in Fig. 6, the roll, pitch and yaw constant
momentum
mappings do not intersect at a single point, which indicates that for the
selected thruster and
pre-determined burn time, an exact angular momentum unloading operation is not
possible.
For example, setting the thruster angles to those of point A would correct
properly for the roll
and yaw, but not for the pitch. The analysis module 54 of Fig. 4 can be used
to find a
compromise solution for the orientation of the thruster such that it will
yield an angular
momentum unloading operation compatible with desired angular momentum
corrections and
possibly with other considerations regarding the satellite 20.
A first approach to finding such a compromise solution is to have the analysis
module
54 provide the mappings for each desired roll, pitch and yaw angular momentum
correction
to a display module 56. The display module 56 is in communication with a
pointing device
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58, such as a mouse, in communication with readout module 60. A graphics
module included
in the display module 56 is programmed to display an interface 66 with a graph
68 of the
mappings as shown in Fig. 7. The interface 66 includes a cursor 62, controlled
by the
pointing device 58, and a cursor readout 63 controlled by readout module 60
displaying the
coordinates of the position occupied by the cursor 62. Additionally, the
calculation module
50 can provide the torque mappings as a function of the angles p and y to the
analysis module
54 and the interface 66 can include an angular momentum readout 64 associated
with the
position of the cursor 62 for displaying the roll, pitch and yaw angular
momentum values, i.e.
the values of the torque mappings multiplied by the bum time for the
coordinates of the
cursor 62. The cursor readout 63 and the angular momentum readout 64 can be
real-time
readouts, i.e. readouts that are updated substantially at the same time as the
cursor is moved
on the graph 68, or can be updated in non-real-time.
Thus, an operator can easily determine the angular momentum unloading
parameters
of the satellite 20 using a selected thruster 22. To do this, the operator
simply inputs the
desired angular momentum corrections, the thruster ID and the desired bum time
to the
calculation module 50. The analysis module 54 receives the above-mentioned
mappings from
the calculation module 50 and displays the constant angular momentum mappings
in the
graph 68 of the interface 66. By using the pointing device 58 to position the
cursor 62 in an
area of the graph 68 where the three constant angular momentum plots intersect
each other,
the operator can determine the thruster angular coordinates values from the
cursor readout 63
and the corresponding angular momentum corrections from the readout 64.
If the operator is satisfied that the thruster angular coordinates and their
associated
angular momentum corrections are adequate, the thruster ID, bum time and
angular
coordinates of the thruster can be uploaded to the spacecraft 20 by the
satellite command
center 41 as shown in Fig. 3. These parameters are received by the transceiver
40 of the
satellite 20 and the thruster controller 38, in accordance with the uploaded
parameters, effects
an angular momentum unloading operation.
Alternatively, instead of selecting a thruster 22 to unload desired angular
momentum
values from the satellite for a given thruster bum time, determining the
angles of the selected
thruster and effecting the angular momentum unloading operation in accordance
with the
determined angles, an operator can determine the angles for each thruster 22
for a given burn
time and select the thruster yielding the best result. For example, in the
case of a satellite 20
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having four thrusters 22, the display module 56 can display four interfaces
similar to interface
66 with one thruster per interface. The operator can then look at the graphs
68 of each display
and decide which one depicts an optimum angular momentum unloading operation.
In most
cases, it will be the one where the closed surface 70, shown in Fig. 7, is the
smallest. The
operator can repeat this step for a plurality of thruster burn time and select
the best thruster
based on the plurality of plots generated by the system of Fig. 4. Ultimately,
the satellite
operator uploads the thruster ID, burn time and thruster orientation values to
the satellite 20.
Another approach for determining the angular momentum unloading operation
parameters is described in relation to Fig. 8 where the closed surface 70 and
the intersection
points A, B, and C of the constant angular roll, pitch and yaw momentums
mappings are
depicted. As will be understood by a skilled worker in the art, the analysis
module can be
programmed to calculate the centroid 72 of the closed surface 70 and a
centroid readout 74
can display the centroid coordinates along with the corresponding angular
momentum
corrections. Alternately, only the centroid coordinates are displayed in
readout 74.
Yet another approach in determining the angular momentum unloading operation
parameters is described in relation to Fig. 9 where the closed surface 70 and
the intersection
points A, B, and C of the constant angular roll, pitch and yaw momentums
mappings are
depicted. As will be understood by a skilled worker in the art, the analysis
module can be
programmed to determine a point 76 where a root mean square error of angular
momentum
correction is minimized. The optimized coordinates of point 76 together with
the
corresponding angular momentum corrections can be displayed in readout 78.
Fig. 10 depicts a flow chart a method according to a non-limiting embodiment
of the
present invention. At step 80, the desired angular momentum corrections are
determined. At
step 82, for a single thruster and pre-determined burn time, a constant
angular momentum
mapping of values of a first thruster orientation parameter to values of a
second thruster
orientation parameter is generated for each desired angular momentum
correction. At step 84,
the orientation of the thruster is selected in accordance with the mappings.
The selection of a particular thruster 22 and thruster burn time to effect a
desired
angular momentum correction can be based on many factors. For example, the
thruster, and
its respective burn time, can be chosen to produce as little drift as
possible, or can be chosen
for operational reasons, such as fuel availability. The quantification of what
is low drift takes
into account the operational box of the satellite, i.e. the ranges of the
satellite's parameters
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that must conformed to for proper operation of the satellite and, the
periodicity of momentum
unloading and station-keeping operations. For example, a maneuver resulting in
relatively
large drift can be allowed if it is known that a subsequent maneuver to occur
shortly will
correct for this drift and that the satellite will remain within its
operational box for the time
being.
As will be understood by a worker skilled in the art, generating the torque
and
constant angular momentum mappings mentioned above can be accomplished
relatively
straightforwardly through various reference frame transfer matrices. For
example, the
reference frame transfer matrix to go from the thruster reference frame to the
gimbaled
platform having the thruster can be expressed as:
Cos(-p) Sin(-y)Sin(-p) -Cos(- y)Sin(- p)
0 Cos(- y) Sin(-y)
Sin(- p) -Sin(- y) Cos(-p) Cos(- y)Cos(- p)
where y and p are orientation parameters of the thrusters.
The transfer matrix to move from the gimbaled platform to the reference frame
of the
satellite body is simply a constant matrix i.e. independent of any orientation
of the thruster
gimbals. The above matrices operate on the center of mass vector pointing from
the center of
mass of the satellite to the gimbaled XIPS platform (GXP), on the vector
following the GXP
and the vectors associated with the XIPS lever arms. Thus, a thruster's thrust
vector can be
resolved in the reference frame of satellite body using the above transfer
matrices.
The embodiments of the present invention practiced with a single Xenon Ion
Propulsion System (XIPS) thruster, such as the Boeing Aerospace 702 thrusters,
was found to
require a very short thruster burn time in comparison with known two-XIPS-
thruster angular
momentum unloading operations. For example, a two-XIPS-thruster angular
momentum
unloading operation requiring 4 hours can be done with the one thruster
approach according
to embodiments of the present invention in as little as 10 minutes. The
resulting decrease in
electricity consumption and the reduced allocation of time and resources to
the angular
momentum operation are significant.
In the preceding description, for purposes of explanation, numerous details
were set
forth in order to provide a thorough understanding of the present invention.
However, it will
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be apparent to one skilled in the art that these specific details are not
required in order to
practice the embodiments of the present invention. In other instances, well-
known electrical
structures and circuits were shown in block diagram form in order not to
obscure the present
invention. For example, specific details are not provided as to whether the
embodiments of
the invention described herein are implemented as a software routine, hardware
circuit,
firmware, or a combination thereof.
Embodiments of the present invention may be represented as a software product
stored on a machine-readable medium (also referred to as a computer-readable
medium, a
processor-readable medium, or a computer usable medium having a computer
readable
program code embodied therein). The machine-readable medium may be any type of
magnetic, optical, or electrical storage medium including a diskette, compact
disk read only
memory (CD-ROM), memory device (volatile or non-volatile), or similar storage
mechanism.
The machine-readable medium may contain various sets of instructions, code
sequences,
configuration information, or other data which, when executed, cause a
processor to perform
steps in a method according to an embodiment of the invention. Those of
ordinary skill in the
art will appreciate that other instructions and operations necessary to
implement the described
invention may also be stored on the machine-readable medium. Software running
from the
machine-readable medium may interface with circuitry to perform the described
tasks.
Thus, embodiments of the present invention provided a method and system for
performing a single thruster angular momentum unloading operation of a
satellite. The single
thruster approach is greatly beneficial over two-thruster approaches in that
the user has
greater flexibility to define the thruster bum time to use, and the required
thruster bum time is
significantly reduced thereby decreasing the amount of energy required in the
angular
momentum unloading operation. This is particularly important when the
thrusters used in the
unloading operation are thrusters operating on electricity, such thrusters
being, e.g., XIPS
thrusters.
The above-described embodiments of the present invention are intended to be
examples only. Alterations, modifications and variations may be effected to
the particular
embodiments by those of skill in the art without departing from the scope of
the invention,
which is defined solely by the claims appended hereto.
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