WO2007128112A1

WO2007128112A1 - System and method for unloading angular momentum from a spacecraft momentum wheel stabilization system

Info

Publication number: WO2007128112A1
Application number: PCT/CA2007/000774
Authority: WO
Inventors: John B. Holland
Original assignee: Bce Inc.
Priority date: 2006-05-08
Filing date: 2007-05-04
Publication date: 2007-11-15
Also published as: CA2546201A1

Abstract

A system and method for determining an orientation of a thruster used in unloading angular momentum from a spacecraft. The system and method use only one thruster in the unloading operation. The system and method allow for user flexibility in choosing maneuver duration to complement on-board maneuver plans or to avoid possible conflicts with operations limitations. When used with XIPS thrusters, the required thruster burn time for a single thruster unloading operation is greatly reduced with respect to known two-XIPS-thruster angular momentum unloading operations. A graphical user interface can be used by a satellite operator to determine the specifics of the angular momentum unloading operation.

Description

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 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 burn 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^th 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 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 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 burn 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 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 bum 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-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. 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.

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 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. 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 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. 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, 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 (γ).

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 burn 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 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 γ 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 burn 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 ED and the desired burn 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, burn 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 burn 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 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 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(-γ)Sin(-p) -Cos(- γ)Sin(- p)

0 Cos(- Y) Sin(-γ)

Sin(- p) -Sin(- Y) Cos(-p) Cos(- γ)Cos(- p)

where γ 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 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 burn time to use, and the required thruster burn 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.

Claims

CLAIMS:

1. A method of unloading angular momentum from a spacecraft based on desired angular momentum corrections, each desired momentum correction associated with a spacecraft momentum axis, the method comprising: 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; and selecting a specific orientation of the thruster in accordance with the mappings.

2. The method of claim 1, further comprising communicating the specific orientation of the thruster to the spacecraft.

3. The method of claim 2, further comprising directing the thruster in accordance with the specific orientation.

4. The method of claim 1, further comprising firing the thruster for the desired thruster burn time.

5. The method of claim 1 , wherein generating the constant angular momentum mapping of values of the first orientation parameter to values of the second orientation parameter is repeated for a plurality of thruster burn times.

6. The method of any one of claims 1 to 5, wherein selecting the specific orientation of the thruster comprises determining values of first and second orientation parameters for which an error in angular momentum correction is minimized.

7. The method of any one of claims 1 to 6, wherein selecting the specific orientation of the thruster further comprises displaying the mappings on a graph of the first orientation parameter as a function of the second orientation parameter.

8. The method of any one of claims 1 to 7, wherein selecting the specific orientation of the thruster further comprises determining intersection points of the mappings for at least one thruster burn time.

9. The method of claim 8, wherein selecting the specific orientation of the thruster further comprises determining a surface area of a closed shape defined by the intersection points and by the constant angular momentum mappings.

10. The method of claim 9, wherein selecting the specific orientation of the thruster further comprises determining a centroid of the closed shape.

11. The method of claim 7, wherein selecting the specific orientation of the thruster further comprises moving a cursor on the graph, the cursor being associated with a readout for displaying coordinates of a position of the cursor on the graph.

12. The method of any one of claims 1 to 11, wherein the first orientation parameter is a first angle p and the second orientation parameter is a second angle γ.

13. The method of claim 12, wherein the angles p and γ are related to gimbaled arms to which the thruster is attached.

14. The method of claim 6, wherein the error in angular momentum correction is a root mean square error.

15. The method of any one of claims 1 to 14, wherein selecting a specific orientation of the thruster comprises selecting the specific orientation of the thruster in accordance with operational limits of the first and second orientation parameters.

16. The method of any one of claims 1 to 15, wherein determining desired angular momentum corrections comprises determining desired angular momentum corrections in accordance with a history of angular momentum corrections.

17. The method of any one of claims 1 to 15, wherein determining desired angular momentum corrections comprises determining desired angular momentum corrections in accordance with a pre-determined schedule of angular momentum corrections.

18. The method of any one of claims 1 to 15, wherein determining desired angular momentum corrections comprises receiving telemetry data from the spacecraft.

19. The method of claim 18, wherein the telemetry data includes data of a momentum wheel stabilization system of the spacecraft.

20. The method of claim 19, wherein, the desired angular momentum corrections are the negative of angular momentum values of the momentum wheel stabilization system.

21. The method of claim 19, wherein the desired angular momentum corrections are biased angular momentum values of the momentum wheel stabilization system.

22. A system for determining an orientation of a thruster used in unloading angular momentum from a spacecraft, the system comprising: a calculation module 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; and an analysis module operable to analyze the mappings.

23. The system of claim 22, further comprising an input module operable to input the desired momentum corrections.

24. The system of claim 22 or 23, further comprising an input module operable to input the at least one pre-determined burn time.

25. The system of any one of claims 22 to 24, further comprising: a graph module operable to generate a plot of the first orientation parameter as a function of the second orientation parameter for each mapping; and a display module operable to display the plot.

26. The system of any one of claims 22 to 25, wherein the analysis module comprises an optimization module operable to determine first and second orientation parameter values that minimize an error in angular momentum correction.

27. The system of claim 26, wherein the error in angular momentum correction is a root mean square error.

28. The system of any one of claims 22 to 25, wherein the analysis module comprises a centroid module operable to determine a centroid of a closed surface defined by the mappings and intersections points of mappings of a same pre-determined burn time.

29. The system of any one of claims 25 to 28, further comprising: a pointing device operable to move a cursor on the graph; and a real-time cursor readout module operable to display coordinates of a position of the cursor on the graph.

30. The system of claim 29, wherein the coordinates are angles of the thruster.

31. The system of claim 29 or claim 30, wherein the real-time cursor readout module is further operable to display angular momentum values associated to the position of the cursor on the graph.

32. A method for selecting a single thruster for an angular momentum unloading operation from a spacecraft, the method comprising: 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 burn time; and selecting one of the at least two thrusters in accordance with the constant angular momentum mappings of the at least two thrusters.

33. A system for selecting a thruster for an angular momentum unloading operation from a spacecraft, the system comprising: a calculation module operable to generate, in accordance with at least one predetermined 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; and an analysis module operable to analyze the mappings.

34. The system of claim 33, further comprising an input module operable to input the desired momentum corrections.

35. The system of claim 33 or 34, further comprising an input module operable to input the at least one pre-determined burn time.

36. The system of any one of claims 33 to 35, further comprising: a graph module operable to generate a graph of the first orientation parameter as a function of the second orientation parameter for each mapping for each of the at least two thrusters; and a display module operable to display the graphs.

37. The system of any one of claims 33 to 36, wherein the analysis module comprises an optimization module operable to determine first and second orientation parameter values that minimize an error in angular momentum correction for each thruster.

38. The system of claim 37, wherein the error in angular momentum correction is a root mean square error.

39. The system of any one of claims 33 to 36, wherein the analysis module comprises a centroid module operable to determine centroids of closed surfaces defined by the mappings and intersections points of mappings of a same pre-determined burn time.

40. The system of claim 36, further comprising: a pointing device operable to move a cursor on the graphs; and a real-time cursor readout module operable to display coordinates of a position of the cursor on the graphs.

41. The system of claim 40, wherein the coordinates are angles of the thruster.

42. The system of claim 40 or claim 41, wherein the real-time cursor readout module is further operable to display angular momentum values associated to the position of the cursor on the graph.

43. A system for unloading angular momentum from a spacecraft, the system comprising: a gimbaled thruster attached to the spacecraft and positionable relative thereto; a sub-system operable to determine an orientation of the gimbaled thruster, the subsystem comprising: a calculation module 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; and an analysis module operable to analyze the mappings to determine an orientation of the gimbaled thruster to unload desired angular momentum from the spacecraft; and a thruster controller operable to receive the orientation of the gimbaled thruster from the sub-system.

44. The system of claim 43, wherein the thruster controller is further operable to receive a pre-determined thruster burn time.

45. The system of claim 44, wherein the thruster controller is further operable to control the gimballed thruster in accordance with the received orientation and the predetermined thruster burn time.

46. A system for visualizing orientation parameters of a thruster used in an angular momentum unloading operation of a spacecraft, the system comprising: a calculation module operable to generate, in accordance with at least one predetermined 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; an analysis module operable to analyze the mappings. a graph module operable to generate a graph of the first thruster parameter as a function of the second thruster parameter for each mapping; and a display module operable to display the graph.

47. The system of claim 46, further comprising an input module operable to input of desired momentum corrections.

48. The system of claim 46 or claim 47 further comprising an input module operable to input the at least one pre-determined burn time.

49. The system of any one of claims 46 to 48, further comprising: a pointing device operable to move a cursor on the graph; and a real-time cursor readout module operable to display coordinates of a position of the cursor on the graph.

50. The system of claim 49, wherein the coordinates are angles of the thruster.

51. The system of claim 49 or claim 50, wherein the real-time cursor readout module is further operable to display angular momentum values associated with the position of the cursor on the graph.

52. An apparatus for unloading angular momentum from a spacecraft based on desired angular momentum corrections, each desired momentum correction associated with a spacecraft momentum axis, the apparatus comprising: 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.