Optimal sun safe attitude for satellite ground tracking

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.

BACKGROUND OF THE INVENTION

1. Field of the Invention

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.

2. Description of the Related Art

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.

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's attitude control subsystems. However, some satellite missions demand the use of sensor packages that substantially increase the satellite's moment of inertia about at least one of the satellite'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'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.

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

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 IXXand about the second axis IYYand a moment of inertia IZZabout a third (z) axis perpendicular to the first axis (x) and the second (y) axis such that IXX<IYYand IXX<IZZ. 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'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.

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.

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'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.

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.

DETAILED DESCRIPTION OF PREFERRED EMBODIMENTS

FIG. 1illustrates a three-axis stabilized satellite or spacecraft100. The spacecraft100is preferably situated in a stationary orbit about the Earth. The satellite100has a main body102, a pair of solar panels104, a pair of high gain narrow beam antennas106, and a telemetry and command omnidirectional antenna108which is aimed at a control ground station. The satellite100may also include one or more sensors110to measure the attitude of the satellite100. 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 inFIG. 1are referred to by the numerals104N and104S for the “North” and “South” solar panels, respectively. In the illustrated embodiment, the solar panels104can 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.

The three axes of the spacecraft10are shown inFIG. 1. The pitch axis P lies along the plane of the solar panels140N and140S. The roll axis R and yaw axis Y are perpendicular to the pitch axis P and lie in the directions and planes shown.

The satellite100also comprises a payload such as a sensor system having a sensor array112. In the illustrated embodiment, the sensor array112implements a space-based radar system used to view targets on the Earth. The sensor array112could be coupled to the satellite bus102via one or more joints and motors that would permit the sensor array112to be rotated along it's longitudinal (x) axis to view the Earth, or may be fixedly coupled to the satellite bus102so that any reorientation of the sensor array112requires that the satellite bus be reoriented as well. In the illustrated embodiment, the z axis is perpendicular to the sensor array.

Accordingly, in the illustrated embodiment, the satellite has a sensor array112disposed longitudinally along the x axis and rotatable solar panels104disposed longitudinally along a second y axis perpendicular to the x axis. The satellite100has a moment of inertia IXXabout the x axis, a moment of inertia IYYabout the y axis and a moment of inertia IZZabout the z axis perpendicular to the first axis (x) and the second (y) axis. Also in the illustrated embodiment, the mass distribution of the satellite100is 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., IXX<IYYand IXX<IZZ).

Due to the length and corresponding mass properties of the sensor array112and 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 plane402. 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.

FIG. 2is a diagram depicting the functional architecture of a representative attitude control system200. Control of the satellite100is provided by a computer or spacecraft control processor (SCP)202. The SCP202performs 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.

Input to the spacecraft control processor202may come from any combination of a number of spacecraft components and subsystems, such as a transfer orbit sun sensor204, an acquisition sun sensor206, an inertial reference unit208, a transfer orbit Earth sensor210, an operational orbit Earth sensor212, a normal mode wide angle sun sensor214, a magnetometer216, and one or more star sensors218.

The SCP202generates control signal commands220which are directed to a command decoder unit222. The command decoder unit operates the load shedding and battery charging systems224. The command decoder unit also sends signals to the magnetic torque control unit (MTCU)226and the torque coil228.

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 thrusters236and momentum wheels242and244of a number and orientation to permit the satellite100to be rotated about any one or all of the x, y, and z axes. Other attitude control actuators may also be used.

The SCP202also sends control commands230to the thruster valve driver unit232which in turn controls the liquid apogee motor (LAM) thrusters234and the attitude control thrusters236.

Wheel torque commands262are generated by the SCP202and are communicated to the wheel speed electronics238and240. These effect changes in the wheel speeds for wheels in momentum wheel assemblies242and244, respectively. The speed of the wheels is also measured and fed back to the SCP202by feedback control signal264.

The spacecraft control processor also sends jackscrew drive signals266to the momentum wheel assemblies243and244. 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 signal268to the spacecraft control processor. The signals268are also sent to the telemetry encoder unit258and in turn to the ground station260.

The spacecraft control processor also sends command signals254to the telemetry encoder unit258which in turn sends feedback signals256to the SCP202. This feedback loop, as with the other feedback loops to the SCP202described earlier, assist in the overall control of the spacecraft. The SCP202communicates with the telemetry encoder unit258, which receives the signals from various spacecraft components and subsystems indicating current operating conditions, and then relays them to the ground station260.

The wheel drive electronics238,240receive signals from the SCP202and control the rotational speed of the momentum wheels. The jackscrew drive signals266adjust 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.

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.

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.

The SCP202may include or have access to memory270, such as a random access memory (RAM). Generally, the SCP202operates under control of an operating system272stored in the memory270, and interfaces with the other system components to accept inputs and generate outputs, including commands. Applications running in the SCP202access and manipulate data stored in the memory270. The spacecraft10may 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.

In one embodiment, instructions implementing the operating system272, 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 system272and the computer program are comprised of instructions which, when read and executed by the SCP202, causes the spacecraft processor202to perform the steps necessary to implement and/or use the present invention. Computer program and/or operating instructions may also be tangibly embodied in memory270and/or data communications devices (e.g. other devices in the spacecraft10or 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.

FIG. 3is 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 satellite100is in a solar energy collection orientation, and in the second (data collection) mode of operation, the satellite100is in a service orientation wherein the sensor array112is directed to obtain data from terrestrial objects.

FIG. 3will be discussed in further reference toFIGS. 4-7.FIGS. 4-5illustrate the satellite100in the solar collection mode, whileFIGS. 6-7illustrate the satellite100in the data collection mode.

If the satellite100is not in the sensor data collection mode, logic is routed to block304, where operations related to solar collection are described. If the satellite100is not in the data collection mode, the satellite100is oriented to align the x axis normal to the plane of the satellite's orbit402around the Earth404, as shown in block304and inFIG. 4. In one embodiment, this is accomplished by orienting the satellite100to align the x axis to be perpendicular to the Earth nadir (N)502and the satellite velocity vector (V)504.

To direct the solar panels in the direction of the Sun406, the satellite100is rotated about the x axis by an angle ψ so that the y axis is perpendicular to the line of sight to the Sun403, as shown in block306andFIG. 5. In the illustrated embodiment, this is accomplished by rotating the satellite bus102about the x axis, but this may be accomplished by rotating the solar panels104as well. The solar panels104may also rotated about the y axis by an angle θ from the satellite orbital plane402, as shown in block308andFIG. 4to further orient the solar panels to be perpendicular to the line of sight to the Sun406. This can accomplished, for example by using motors or actuators in the satellite bus102that rotate the solar panels104about the y axis relative to the satellite bus102.

While in the solar collection mode, the satellite100orientation 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 block310. This may be accomplished by rotating the x axis of satellite100about the same axis and at the same rate as nodal precession of the orbital plane402. The satellite100can also be rotated about the x axis to maintain the surface of the solar panels104perpendicular to the Sun line of sight403. The required rotation rate is approximately the reciprocal of the orbital period of the satellite100.

If the satellite100is in the sensor data collection or service mode, the satellite100is rotated only about the first (x) axis to direct the sensitive axis of the sensor array112at one or more terrestrial targets602, as shown in block312and illustrated inFIGS. 6 and 7. Sensor data is then collected, as illustrated in block314. After doing so, the surface of the solar panels104will no longer be perpendicular to the Sun line of sight403, so the amount of solar energy collected by the solar panels will decrease. However, after block302determines the required sensor data is collected, the satellite100returns to the solar collection mode by following steps304-310above.

CONCLUSION

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.

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.