Abstract:
A method and apparatus for maneuvering a satellite in orbit to alternately optimize the collection of solar energy and to take sensor data of terrestrial objects is disclosed The longitudinal axis of a large payload package is oriented perpendicular to the orbital plane to minimize the disturbance torque due to gravity gradient, and to allow simple rotation about the axis for attitude change between optimal Sun and optimal ground coverage.

Description:
BACKGROUND OF THE INVENTION 
       [0001]    1. Field of the Invention 
         [0002]    The present invention relates generally to systems and methods for maneuvering satellites in orbit, and in particular, to a system and method for maneuvering a satellite to direct a one sensors at terrestrially based locations and for optimal collection of solar energy. 
         [0003]    2. Description of the Related Art 
         [0004]    Satellites typically comprise solar panels that are used to collect solar energy. The solar energy capabilities of such panes are maximized if the panes face directly into the Sun (that is, a vector normal to the plane of the solar panes is directed at the Sun). While solar panels are typically capable of rotating along one axis, the satellite itself usually also needs to be oriented to absorb maximum energy from the Sun. 
         [0005]    In most circumstances, this poses no significant problems, as the moment of inertia of the satellite is small enough to permit such satellite orientations without undue requirements placed upon the satellite&#39;s attitude control subsystems. However, some satellite missions demand the use of sensor packages that substantially increase the satellite&#39;s moment of inertia about at least one of the satellite&#39;s axes. For example, sensor arrays used for space-based ground surveillance using radar can be quite large compared to other satellite structures, substantially increasing the satellite&#39;s moment of inertia. Such arrays may also be fixed to the satellite bus itself, thereby requiting the satellite bus to rotate to direct the arrays at the desired terrestrial location. As a result, such satellites can be asked to perform frequent and large rotations about axes with large moments of inertia. This places costly requirements on the satellite attitude control system, particularly the motive elements (e.g. thrusters, momentum wheels and the like) that are used to rotate the satellite. 
         [0006]    What is needed is a system and method that permits satellites with large or extended payloads to achieve their mission while still permitting maximum solar energy absorption without the use of high capacity motive elements. The present invention satisfies this need. 
       SUMMARY OF THE INVENTION 
       [0007]    To address the requirements described above, the present invention discloses a method and apparatus for maneuvering a satellite in orbit to alternately optimize the collection of solar energy and to take sensor data of terrestrial objects. The satellite has a payload such as a sensor disposed longitudinally along a first (x) axis and rotatable solar panels disposed longitudinally along a second (y) axis perpendicular to the first axis (x), the satellite having a moment of inertia about the first axis I XX  and about the second axis I YY  and a moment of inertia I zz  about a third (z) axis perpendicular to the first axis (x) and the second (y) axis such that I XX &lt;I YY  and I XX &lt;I ZZ . When the satellite is not in a payload data collection mode, the satellite is oriented to align the first (x) axis normal to the plane of the orbit and rotated about the first (x) axis such that the second (y) axis is perpendicular to a Sun line of sight. The satellite&#39;s solar panels are also rotated about the second (y) axis at a first angle (θ) from the satellite orbital plane to orient the solar panels towards the Sun. The alignment of the first (x) axis is maintained normal to the satellite orbital plane during the non-payload data collection mode. When the satellite is in a payload data collection mode, the satellite is rotated about only the first (x) axis to direct the payload at a terrestrial target; and target-related payload data is collected. 
         [0008]    The apparatus comprises a plurality of attitude motive elements, the plurality of attitude motive elements for changing the attitude of the satellite in the first (x) axis, the second (y) axis, and the third (z) axis and one or more processors, communicatively coupled to the attitude motive elements. The processor may be communicatively coupled to a memory storing instructions comprising instructions for maneuvering the satellite to a first orientation by orienting the satellite to align the first (x) axis normal to the plane of the orbit, rotating the satellite about the first (x) axis such that the second (y) axis is perpendicular to a Sun line of sight, and rotating the solar panels about the second (y) axis at a first angle (θ) from the satellite orbital plane to orient the solar panels towards the Sun, and maintaining the alignment of the first (x) axis normal to the satellite orbital plane during the non-payload data collection mode when the satellite is not in a payload data collection mode. The memory may also store instructions for maneuvering the satellite to a second orientation by rotating the satellite only about the first (x) axis to direct the payload at a terrestrial target and to collect the payload data when the satellite is in a payload data collection mode. 
         [0009]    For the most satellites with global surveillance mission (including those in low-earth or mid-earth orbits), the motion of the Sun and the precession of the satellite&#39;s orbital plane are slow compared to their relative motion to the ground. Therefore, once the attitude of the satellite for optimal solar power is set, it will require very little or no adjustment over a small number of orbit revolutions. In typical daily operations, the spacecraft will be in the Sun optimal attitude most of the time. The attitude may then be changed only to allow optimal payload sensor coverage when the satellite is near the ground area of interest. 
         [0010]    The design described herein places the rotation axis of the solar panel on the y axis of the bus, and the payload sensor array along the x axis of the bus. Due to the length of the array support structure and the corresponding mass property of the system, the bus x axis is the principal axis with minimum moment of inertia. This design places the x-axis perpendicular to the orbital plane to minimize the disturbance torque due to gravity gradient, and to allow simple rotation about the X-axis for attitude change between optimal Sun and optimal ground coverage. Small rotations at constant rate about an inertial reference can be introduced to accommodate the precession of the orbital plane due to natural perturbation. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS  
         [0011]    Referring now to the drawings in which like reference numbers represent corresponding parts throughout: 
           [0012]      FIG. 1  is a diagram of a representative satellite having a large payload array; 
           [0013]      FIG. 2  is a block diagram illustrating a satellite attitude control system; 
           [0014]      FIG. 3  is a flow chart illustrating representative method steps that can be used to practice one embodiment of the present invention; 
           [0015]      FIG. 4  is a diagram of a satellite in orbit around the Earth in a power charging configuration, viewed edge-on to its orbital plane; 
           [0016]      FIG. 5  is a diagram of the satellite shown in  FIG. 4 , viewed from a location perpendicular to the orbital plane; 
           [0017]      FIG. 6  is a diagram of a satellite in orbit around the Earth in a service orientation, viewed edge-on to the orbital plane; and 
           [0018]      FIG. 7  is a diagram of the satellite shown in  FIG. 6 , viewed from a location perpendicular to the orbital plane. 
       
    
    
     DETAILED DESCRIPTION OF PREFERRED EMBODIMENTS 
       [0019]    In the following description, reference is made to the accompanying drawings which form a part hereof, and which is shown, by way of illustration, several embodiments of the present invention. It is understood that other embodiments may be utilized and structural changes may be made without departing from the scope of the present invention. 
         [0020]      FIG. 1  illustrates a three-axis stabilized satellite or spacecraft  100 . The spacecraft  100  is preferably situated in a stationary orbit about the Earth. The satellite  100  has a main body  102 , a pair of solar panels  104 , a pair of high gain narrow beam antennas  106 , and a telemetry and command omnidirectional antenna  108  which is aimed at a control ground station. The satellite  100  may also include one or more sensors  110  to measure the attitude of the satellite  100 . These sensors may include sun sensors, earth sensors, and star sensors. Since the solar panels are often referred to by the designations “North” and “South”, the solar panels in  FIG. 1  are referred to by the numerals  104 N and  104 S for the “North” and “South” solar panels, respectively. In the illustrated embodiment, the solar panels  104  can be rotated along their longitudinal axis (the y axis) to direct them at the Sun. Other embodiments may include solar panels that are also rotatable about the z and y axes, but such implementations are more expensive, and in light of the invention described below, unnecessary for purposes of maximizing solar power collection. 
         [0021]    The three axes of the spacecraft  10  are shown in  FIG. 1 . The pitch axis P lies along the plane of the solar panels  140 N and  140 S. The roll axis R and yaw axis Y are perpendicular to the pitch axis P and lie in the directions and planes shown. 
         [0022]    The satellite  100  also comprises a payload such as a sensor system having a sensor array  112 . In the illustrated embodiment, the sensor array  112  implements a space-based radar system used to view targets on the Earth. The sensor array  112  could be coupled to the satellite bus  102  via one or more joints and motors that would permit the sensor array  112  to be rotated along it&#39;s longitudinal (x) axis to view the Earth, or may be fixedly coupled to the satellite bus  102  so that any reorientation of the sensor array  112  requires that the satellite bus be reoriented as well. In the illustrated embodiment, the z axis is perpendicular to the sensor array. 
         [0023]    Accordingly, in the illustrated embodiment, the satellite has a sensor array  112  disposed longitudinally along the x axis and rotatable solar panels  104  disposed longitudinally along a second y axis perpendicular to the x axis. The satellite  100  has a moment of inertia I XX  about the x axis, a moment of inertia I YY  about the y axis and a moment of inertia I ZZ  about the z axis perpendicular to the first axis (x) and the second (y) axis. Also in the illustrated embodiment, the mass distribution of the satellite  100  is such that the moment of inertia about the x axis is less than the moment of inertia of either the y or z axes (e.g., I XX &lt;I YY  and I XX &lt;I ZZ ). 
         [0024]    Due to the length and corresponding mass properties of the sensor array  112  and support structure the x axis is the axis with the minimum moment of inertia. As described below, this invention advantageously places the x axis perpendicular to the orbital plane  402 . This results in two important benefits. First, it minimizes disturbance torques due to gravity gradients and second, it allows the satellite to change modes from solar collection to data collection and back again through a simple rotation about only the x axis. 
         [0025]      FIG. 2  is a diagram depicting the functional architecture of a representative attitude control system  200 . Control of the satellite  100  is provided by a computer or spacecraft control processor (SCP)  202 . The SCP  202  performs a number of functions which may include post ejection sequencing, transfer orbit processing, acquisition control, stationkeeping control, normal mode control, mechanisms control, fault protection, and spacecraft systems support, among others. The post ejection sequencing could include initializing to assent mode and thruster active nutation control (TANC). The transfer orbit processing could include attitude data processing, thruster pulse firing, perigee assist maneuvers, and liquid apogee motor (LAM) thruster firing. The acquisition control could include idle mode sequencing, sun search/acquisition, and Earth search/acquisition. The stationkeeping control could include auto mode sequencing, gyro calibration, stationkeeping attitude control and transition to normal. The normal mode control could include attitude estimation, attitude and solar array steering, momentum bias control, magnetic torquing, and thruster momentum dumping (H-dumping). The mechanisms mode control could include solar panel control and reflector positioning control. The spacecraft control systems support could include tracking and command processing, battery charge management and pressure transducer processing. 
         [0026]    Input to the spacecraft control processor  202  may come from any combination of a number of spacecraft components and subsystems, such as a transfer orbit sun sensor  204 , an acquisition sun sensor  206 , an inertial reference unit  208 , a transfer orbit Earth sensor  210 , an operational orbit Earth sensor  212 , a normal mode wide angle sun sensor  214 , a magnetometer  216 , and one or more star sensors  218 . 
         [0027]    The SCP  202  generates control signal commands  220  which are directed to a command decoder unit  222 . The command decoder unit operates the load shedding and battery charging systems  224 . The command decoder unit also sends signals to the magnetic torque control unit (MTCU)  226  and the torque coil  228 . 
         [0028]    The attitude control system comprises a plurality of attitude motive elements or actuators that are used to change the attitude of the satellite in the first (x) axis, the second (y) axis, and the third (z) axis. Such elements include ACS thrusters  236  and momentum wheels  242  and  244  of a number and orientation to permit the satellite  100  to be rotated about any one or all of the x, y, and z axes. Other attitude control actuators may also be used. 
         [0029]    The SCP  202  also sends control commands  230  to the thruster valve driver unit  232  which in turn controls the liquid apogee motor (LAM) thrusters  234  and the attitude control thrusters  236 . 
         [0030]    Wheel torque commands  262  are generated by the SCP  202  and are communicated to the wheel speed electronics  238  and  240 . These effect changes in the wheel speeds for wheels in momentum wheel assemblies  242  and  244 , respectively. The speed of the wheels is also measured and fed back to the SCP  202  by feedback control signal  264 . 
         [0031]    The spacecraft control processor also sends jackscrew drive signals  266  to the momentum wheel assemblies  243  and  244 . These signals control the operation of the jackscrews individually and thus the amount of tilt of the momentum wheels. The position of the jackscrews is then fed back through command signal  268  to the spacecraft control processor. The signals  268  are also sent to the telemetry encoder unit  258  and in turn to the ground station  260 . 
         [0032]    The spacecraft control processor also sends command signals  254  to the telemetry encoder unit  258  which in turn sends feedback signals  256  to the SCP  202 . This feedback loop, as with the other feedback loops to the SCP  202  described earlier, assist in the overall control of the spacecraft. The SCP  202  communicates with the telemetry encoder unit  258 , which receives the signals from various spacecraft components and subsystems indicating current operating conditions, and then relays them to the ground station  260 . 
         [0033]    The wheel drive electronics  238 ,  240  receive signals from the SCP  202  and control the rotational speed of the momentum wheels. The jackscrew drive signals  266  adjust the orientation of the angular momentum vector of the momentum wheels. This accommodates varying degrees of attitude steering agility and accommodates movement of the spacecraft as required. 
         [0034]    The use of reaction wheels or equivalent internal torquers to control a momentum bias stabilized spacecraft allows inversion about yaw of the attitude at will without change to the attitude control. In this sense, the canting of the momentum wheel is entirely equivalent to the use of reaction wheels. 
         [0035]    Other spacecraft employing external torquers, chemical or electric thrusters, magnetic torquers, solar pressure, etc. cannot be inverted without changing the control or reversing the wheel spin direction. This includes momentum bias spacecraft that attempt to maintain the spacecraft body fixed and steer payload elements with payload gimbals. 
         [0036]    The SCP  202  may include or have access to memory  270 , such as a random access memory (RAM). Generally, the SCP  202  operates under control of an operating system  272  stored in the memory  270 , and interfaces with the other system components to accept inputs and generate outputs, including commands. Applications running in the SCP  202  access and manipulate data stored in the memory  270 . The spacecraft  10  may also comprise an external communication device such as a satellite link for communicating with other computers at, for example, a ground station. If necessary, operation instructions for new applications can be uploaded from ground stations. 
         [0037]    In one embodiment, instructions implementing the operating system  272 , application programs, and other modules are tangibly embodied in a computer-readable medium, e.g., data storage device, which could include a RAM, EEPROM, or other memory device. Further, the operating system  272  and the computer program are comprised of instructions which, when read and executed by the SCP  202 , causes the spacecraft processor  202  to perform the steps necessary to implement and/or use the present invention. Computer program and/or operating instructions may also be tangibly embodied in memory  270  and/or data communications devices (e.g. other devices in the spacecraft  10  or on the ground), thereby making a computer program product or article of manufacture according to the invention. As such, the terms “program storage device,” “article of manufacture” and “computer program product” as used herein are intended to encompass a computer program accessible from any computer readable device or media. 
         [0038]      FIG. 3  is a diagram presenting illustrative steps that can be used to practice one embodiment of the present invention. The steps described two modes of operation, each associated with an orientation. In a first (solar collection) mode of operation, the satellite  100  is in a solar energy collection orientation, and in the second (data collection) mode of operation, the satellite  100  is in a service orientation wherein the sensor array  112  is directed to obtain data from terrestrial objects. 
         [0039]      FIG. 3  will be discussed in further reference to  FIGS. 4-7 .  FIGS. 4-5  illustrate the satellite  100  in the solar collection mode, while  FIGS. 6-7  illustrate the satellite  100  in the data collection mode. 
         [0040]    Returning now to  FIG. 3 , block  302  determines whether the satellite  100  is in the sensor data collection mode. If the satellite  100  is in the data processing mode, logic is routed to block  312 , where operations related to sensor data collection mode operations are described. 
         [0041]    If the satellite  100  is not in the sensor data collection mode, logic is routed to block  304 , where operations related to solar collection are described. If the satellite  100  is not in the data collection mode, the satellite  100  is oriented to align the x axis normal to the plane of the satellite&#39;s orbit  402 , as shown in block  304  and in  FIG. 4 . In one embodiment, this is accomplished by orienting the satellite  100  to align the x axis to be perpendicular to the Earth nadir (N)  502  and the satellite velocity vector (V)  504 . 
         [0042]    To direct the solar panels in the direction of the Sun  406 , the satellite  100  is rotated about the x axis by an angle ψ so that the y axis is perpendicular to the line of sight to the Sun  403 , as shown in block  306  and  FIG. 5 . In the illustrated embodiment, this is accomplished by rotating the satellite bus  102  about the x axis, but this may be accomplished by rotating the solar panels  104  as well. The solar panels  104  may also rotated about the y axis by an angle θ from the satellite orbital plane  402 , as shown in block  308  and  FIG. 4  to further orient the solar panels to be perpendicular to the line of sight to the Sun  406 . This can accomplished, for example by using motors or actuators in the satellite bus  102  that rotate the solar panels  104  about the y axis relative to the satellite bus  102 . 
         [0043]    While in the solar collection mode, the satellite  100  orientation in the solar collection mode is maintained by orienting the satellite to maintain the alignment of the x axis to be normal to the satellite orbital plane, as shown in block  310 . This may be accomplished by rotating the x axis of satellite  100  about the same axis and at the same rate as nodal precession of the orbital plane  402 . The satellite  100  can also be rotated about the x axis to maintain the surface of the solar panels  104  perpendicular to the Sun line of sight  403 . The required rotation rate is approximately the reciprocal of the orbital period of the satellite  100 . 
         [0044]    If the satellite  100  is in the sensor data collection or service mode, the satellite  100  is rotated only about the first (x) axis to direct the sensitive axis of the sensor array  112  at one or more terrestrial targets  602 , as shown in block  312  and illustrated in  FIGS. 6 and 7 . Sensor data is then collected, as illustrated in block  314 . After doing so, the surface of the solar panels  104  will no longer be perpendicular to the Sun line of sight  403 , so the amount of solar energy collected by the solar panels will decrease. However, after block  302  determines the required sensor data is collected, the satellite  100  returns to the solar collection mode by following steps  304 - 310  above. 
       CONCLUSION 
       [0045]    This concludes the description of the preferred embodiments of the present invention. The foregoing description of the preferred embodiment of the invention has been presented for the purposes of illustration and description. It is not intended to be exhaustive or to limit the invention to the precise form disclosed. Many modifications and variations are possible in light of the above teaching. For example, while the foregoing was described with respect to a sensor array, the foregoing can be practiced with any payload. Further, although the foregoing is described in terms of being implemented by a processor executing instructions stored in a memory, using the foregoing teaching, it is apparent that the invention may also be implemented by one or more special purpose processors performing subsets of the defined operations, or by hardware dedicated to the tasks. 
         [0046]    It is intended that the scope of the invention be limited not by this detailed description, but rather by the claims appended hereto. The above specification, examples and data provide a complete description of the manufacture and use of the composition of the invention. Since many embodiments of the invention can be made without departing from the spirit and scope of the invention, the invention resides in the claims hereinafter appended.