Method and apparatus for compensating for solar torque transients on a satellite during a solar eclipse

Abstract: A system compensating for solar pressure disturbance on an orbiting satellite comprises sensor means (34) for measuring the satellites attitude (ϑ) relative to a predetermined satellite axis and providing an attitude measurement signal (ϑ m ) in the predetermined axis. Further provided are control means (36) for receiving the attitude measurement signal (ϑ m ) and calculating satellite attitude compensation control signals (T c ) in the predetermined axis. An eclipse transient compensation device (40) provides an eclipse signal to the control means (36) indicating the satellite is being eclipsed from the sun. Said control means (36) removes signals (T i ) being applied to the actuation means (38) from the control means (36) for compensating solar pressure on the satellite. Further are provided actuation means (38) for actuating the satellite in a predetermined axis in response to the compensation control signals (T c ) from the control means (36) (Fig. 2). Further is disclosed a respective method of maintaining a orbiting satellite in a predetermined satellite attitude against the disturbance of solar pressure on the satellite.

Description: BACKGROUND OF THE INVENTION 1. Technical Field This invention relates generally to a method of compensating for pointing of an orbiting satellite, and more particularly, to a method for compensating for the effects of solar pressure on the pointing direction of an orbiting satellite as the satellite goes into and out of eclipse by the sun. 2. Discussion A geosynchronous earth orbit, as is known in the art, is the orbit about the earth in which a satellite or spacecraft will remain fixed above a specific location on the earth. This orbit is at a distance of approximately 35885 Km (22,400 miles) above the earth. In this orbit, a beam, such as a communications beam, from the satellite can be maintained over a desirable area on the earth, such as a particular country, thus establishing an area which will receive the beam. To remain in a geosynchronous orbit it is necessary that the satellite be in an orbit substantially within the equatorial plane of the earth at the desirable distance, and that the satellite's attitude be oriented perpendicular to this plane. Any deviation or disturbance which causes the satellite to direct its antenna away from a boresight location on the earth tends to effect the coverage area of the beam, and thus, produces undesirable results. Many different forces are in effect on the satellite which tend to alter the satellite's antenna pointing direction. As a first order method for countering the effects of the different forces acting on the satellite, it is known to stabilize the satellite's attitude by providing an angular bias momentum which resists changes in the satellite's orientation due to external forces transverse to the bias momentum axis. Satellites using this technique are generally referred to as "momentum bias" satellites. Angular momentum bias is usually provided by a number of momentum or reaction wheels which spin at least part of the satellite. The bias axis set by the spin of the momentum wheels is generally perpendicular to the direction of the orbit of the satellite. Although the bias momentum resists changes in the satellite's orientation in directions transverse to the bias momentum axis, it is still necessary to provide control for correcting variations in the satellite's orientation along the bias axis. Different methods of controlling the satellite's attitude, such as feedback loops, are known in the art. For most bias momentum satellites, the satellite payload, i.e., the part of the satellite carrying at least the antenna, is oriented differently than the momentum wheel. It is therefore necessary to provide means for correcting the orientation of the payload with respect to the orientation of the momentum attitude. Typically, the satellite's payload is defined in three axes referred to as the yaw, roll and pitch axes. If the satellite is in a geosynchronous orbit, the yaw axis is generally directed from the satellite to the center of the earth, the pitch axis is generally directed normal to the plane of the orbit of the satellite and the roll axis is generally perpendicular to the yaw and pitch axes, in a direction of travel of the satellite as is well known in the art. One particular disturbance torque which affects the pointing of a geosynchronous satellite is the disturbance torque which can be equated to solar pressure caused by particles from the sun hitting the satellite. A satellite in a geosynchronous orbit will be eclipsed from the sun one time per revolution of the earth, and thus, the solar pressure on the satellite will go to zero during the eclipse. Typically, the attitude control mechanisms in use on the satellite will compensate for the effect of solar pressure. However, on a momentum bias satellite without direct yaw sensing, disturbance torques are typically estimated with relatively long time constants. These estimated disturbance torques are used by the satellite's attitude control system to create equal and opposite control torques to eliminate or reduce satellite attitude transients. Compared with these time constants, the solar pressure disturbance torque undergoes a rapid transient during eclipses, both when going from full solar pressure disturbance to zero disturbance, and then returning to full disturbance from zero, generally within a matter of minutes. This change is too fast for the disturbance torque sensing to track, and thus leads to potentially large satellite attitude disturbances, especially about the yaw axis. What is needed then is a method to calculate the changes in solar disturbance torques caused by eclipses in order to compensate for this effect on the pointing of the satellite. It is therefore an object of the present invention to provide such a method. SUMMARY OF THE INVENTION This invention discloses a method according to claim 1 determining the presence of a solar eclipse on a satellite, and then compensating for the effect of no solar pressure on the satellite as to the satellite's pointing direction. More particularly, when the solar pressure from the sun goes to zero or is significantly reduced as a result of an eclipse, electrical power delivered by the solar cells also goes to zero and the batteries in the satellite automatically begin discharging to provide the necessary power to operate the satellite. According to one preferred embodiment of the present invention, the battery discharge current is monitored and when it exceeds a predetermined threshold as would be caused by the battery discharging, estimated solar disturbance torques to compensate for solar pressure are set to zero. Once the battery discharge current falls below the threshold, i.e., after the eclipse, the estimated solar disturbance torques are reinstated. It is possible to measure the current indicating solar eclipse by other methods such as current through the solar panels or some other electrical current which changes when the satellite enters and leaves solar eclipses. The invention also relates to a system according to claim 7. Additional objects, advantages, and features of the present invention will become apparent from the following description and appended claims, taken in conjunction with the accompanying drawings. BRIEF DESCRIPTION OF THE DRAWINGS FIG. 1 is a diagram of a momentum bias satellite orbiting the earth which may incorporate compensation for the effects of solar pressure according to a preferred embodiment of the present invention. FIG. 2 is a schematic block diagram depicting a method of compensating for the effects of solar pressure according to one preferred embodiment of the present invention; and FIG. 3 is a schematic block diagram of a second method of compensating for the torque disturbances caused by solar pressure according to a second preferred embodiment of the present invention. DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT The following discussion of the preferred embodiments concerning solar pressure disturbance compensation of a satellite's attitude is merely exemplary in nature and is in no way intended to limit the invention or its application or uses. Before going on to the specifics of the invention, it may be helpful to get an overview of the type of satellite which may incorporate the invention. To get this understanding, first turn to FIG. 1. In FIG. 1, the relationship of a momentum bias satellite 10 orbiting the earth 20 and relative to the sun 22 is shown. Satellite 10 includes a satellite central body 12, here depicted as a cube. By representing body 12 of satellite 10 as a cube it is easier to visualize the coordinates of the satellite's axes of yaw, roll and pitch. As discussed above, the yaw axis is the axis directed from the attitude of the satellite body 12 to the center of the earth 20, the pitch axis is directed normal to the orbit of satellite 10 and normal to the yaw axis, and the roll axis is perpendicular to the yaw and pitch axes, in a direction along travel of satellite 10. Each of these axes are shown in a directional manner in FIG. 1. Satellite 10 includes a momentum wheel 14 which spins in a particular direction, here clockwise as shown, in order to stabilize satellite 10 in its orbit. The axis of rotation of momentum wheel 14 is generally parallel to the pitch axis of the satellite's body 12. The momentum wheel 14 is typically oriented separate from the reference of the satellite body 12. In one example, the orientation of the momentum wheel 14 will be on a 2-axis gimbal. The momentum wheel 14, in this example, is an actuator which adjusts and maintains the satellite's attitude. The pitch axis and momentum axis would be oriented substantially perpendicular to the equatorial plane of the earth 20 if satellite 10 were in a geosynchronous orbit. Satellite 10 further includes an array of integrating gyros 18 and a thermopile earth sensor 24. Integrating gyros are generally used to orient the satellite 10 in a desirable configuration when the satellite 10 reaches its geosynchronous altitude as is known in the art. In normal operation of satellite 10, integrating gyros 18 would not be used. Earth sensor 24, sometimes referred to as a horizon sensor, senses the horizon of the earth in order to establish an inertial reference frame for the proper orientation of the satellite 10. An earth sensor will typically measure the satellite's orientation in the roll and pitch directions. Other types of sensors, including but not limited to star sensors, sun sensors, gyros and ground beacon sensors, could also provide the necessary sensing of the satellite's attitude. Projecting from satellite body 12 are dual solar wings 16 and 17 shown here in a cut-away format. Solar wings 16 and 17 are directed perpendicular to the rays of the sun in order to provide the necessary power to the systems on satellite 10. As discussed above, the solar pressure applied to an orbiting satellite from the sun causes the satellite to deviate from its desirable attitude. Since these satellites are typically equipped with sensing mechanisms in order to determine attitude offsets, the solar pressure generally is compensated for by the attitude control mechanisms included on the satellite. When a satellite goes behind the earth and is eclipsed from the sun, solar pressure goes to zero, generally, within a matter of minutes. Due to the long reaction time of the attitude control compensation algorithms, because the sensors are calibrated to reduce errors, the negation of the solar pressure during an eclipse causes the satellite control mechanisms to overcompensate for the solar pressure and thus the satellite's attitude is affected undesirably. Consequently, this invention proposes detecting the presence of a solar eclipse in order to eliminate the compensation due to solar pressure during the eclipse relatively quickly. In a particular case, where the disturbance torques due to solar pressure cannot be discriminated from other disturbance torques, such as magnetic disturbance, etc., but are the dominant disturbance torques, the entire portion of the control torque which compensates constant disturbance torques could be set to zero. Now turning to FIG. 2, an implementation of a solar pressure compensation measuring scheme is shown in block diagram form according to one preferred embodiment of the present invention. More particularly, a compensation system 30 includes satellite dynamics 32 which represents the satellite on which the disturbance torques T d act upon, including those of solar pressure. The satellite dynamics 32 registers an attitude angle  which is sensed by an attitude sensor 34. In this case, Θ represents any or all of the satellite axes of pitch, roll and yaw. Attitude sensor 34 can be a variety of known attitude sensors, such as an earth sensor, ground beacon sensor, sun sensor or star sensor, as discussed above. Attitude sensor 34 generates an attitude measurement  m and applies it to a summation device as shown. Also applied to the summation device is a control angle signal  c which is the desired attitude of the satellite. An error signal  e is applied to an attitude control compensator 36 as the difference between the control angle signal  c and the measured angle  m . The operation of attitude control compensator 36 will be discussed in greater detail below. The output T c of attitude control compensator 36 is a control torque which is applied to a control actuator 38. Control actuator 38 applies a compensation to the satellite attitude against the disturbance torques of T d . For our satellite 10 above, the actuator is momentum wheel 14. However, other known actuators such as reaction wheels, thrusters, control moment gyros, etc. are also applicable. An eclipse transient compensation device 40 applies an input to attitude control compensator 36 in order to institute or withdraw the compensation control for compensating for solar pressure as will be discussed below. The attitude control compensator 36 includes a proportional feedback compensation system 42, here labeled F(s), and an integrator system 44. The proportional feedback compensation system 42 includes a filtering characteristic which may also include lead/lag compensation as is known in the art. Proportional feedback compensation system 42 applies a gain to the attitude error signal  e to develop a control torque T p . In most cases the value of T p will not be significant enough to completely null the attitude error. Therefore, integrator system 44 is included in order to provide an integrating function, here K I /S, to the gain torque T p in order to completely null the disturbance. In other words, the integrating system 44 continuously integrates T p over time by the function K I /S to build up a control torque T I which when added to the control torque T p will eventually be large enough to null the disturbance torque. The function K I /S is the Laplace transform function of a known integrator. During the times when the solar pressure is not adding to the disturbance torques, i.e., during an eclipse, it is desirable to remove the integration factor T I from the control torque signal T c because the solar pressure is the major disturbance and thus the value T p alone is sufficient. Therefore, switches 50 and 52 are incorporated on both sides of integrator system 44, as shown, such that eclipse transient compensation device 40 can open the switches when it detects the satellite is being eclipsed from the sun. When the switches are open, T I is set to zero or to the known value of non-solar disturbance torques. Eclipse transient compensation device 40 determines when the satellite is being eclipsed from the sun and provides a signal by which the switches 50 and 52 on both sides of integrator system 44 are to be opened during an eclipse. A current measurement device 46 measures the current through a satellite battery circuit (not shown) to determine a time when the batteries are discharging as would occur when the solar wings are not collecting power from the sun. It is noted that the current could be measured in other applicable circuits, such as the solar wing circuits, to determine if the batteries are discharging at a level which would indicate the solar wings are not being powered. This measured current is applied to a threshold logic circuit 48 which determines if the measured current is above a predetermined threshold. If the current traveling through the battery circuit is above a predetermined threshold, it is evidence that an eclipse is occurring and thus, the threshold circuit 48 will apply a signal on its output lines to open the switches 50 and 52 as discussed above. Once the satellite comes out from the eclipse and the solar cells on the solar wings again begin to charge, the battery stops discharging and the current measurement would then drop below the threshold. Thus, a signal from threshold logic circuit 48 would close the switches 50 and 52 again to compensate for the solar pressure in the manner as discussed above. The embodiment discussed above for FIG. 2, specifically concerning the integral action in the attitude control compensation, only has the capability to control disturbance torques which are constant on the satellite. However, solar pressure can also cause disturbance torques which vary sinusoidally in a fixed satellite coordinate system with frequencies at multiples of the orbit rate. More particularly, a "secular" disturbance torque is one which is defined as fixed in inertial space, and thus, oscillates with a frequency of one cycle per orbit of the satellite. These cyclical disturbance torques can be compensated for by the satellite's attitude control system by including notch filters, known to those skilled in the art, at the desired frequencies. Therefore, in order to null these control torques due to the components during an eclipse, the outputs of the notch filters can be nulled during periods when the eclipses are detected. Now turning to FIG. 3, a specific implementation of the attitude control compensator 36 of FIG. 2 is shown in an observer-controller type format where the disturbance torques are explicitly estimated. As before, particular disturbance torgues T d act on the satellite dynamics 32. The satellite dynamics attitude has an angle, here y, which is measured by attitude sensor 34. Attitude sensor 34 generates a measurement signal y m to a satellite attitude observer 54 which processes the measured attitude in a known algorithm depicted here as where: x is the state vector representing the instantaneous state of the satellite dynamics as modeled in the observer; x ∧ is the estimate of x; is the time-derivative of x ∧; u is the control input to the satellite dynamics; d ∧ is the estimate of the disturbance acting on the satellite (i.e., the estimate of T d ); K o x is the observer gain matrix; and A,B,C and L represent the satellite dynamics as modelled in the observer. The resultant value x ∧ is applied to a control algorithm 58 as the estimated attitude state of the satellite. An offset error, here y m - Cx ∧, from the difference between the desirable satellite attitude and the actual satellite attitude is applied to a disturbance torque observer 56 which computes the estimated disturbance torque d ∧ from the offset error by the function d = A d d + K d (y m - C x ) . The value d ∧ is applied to the attitude observer 54 to update the satellite's attitude as well as the control algorithm 58. The control algorithm 58 calculates a control signal u by a known control feedback algorithm u = -K c x - d + K c X c in order to determine the command value u cmd which is given to the control actuator 38 to compensate for the disturbance torques. The control signal u is also applied to the attitude observer 54. As above, a measuring device 46 of eclipse transient compensator 40 measures the current from the discharging batteries on the satellite to determine if an eclipse is present. If the measured current from the discharging battery exceeds a predetermined threshold, then a signal from threshold logic circuit 48 will open a switch 60 in order that the estimated disturbance value d ∧ is not included in the control algorithm 58. When the switch is open, the value of the estimated disturbance torque d ∧ is set to zero or to the known value of non-solar disturbance torques. Consequently, the disturbance torques due to solar pressure is not compensated for during a time when an eclipse is detected. It is noted that the above described systems only have capability of determining two states, whether an eclipse is present or not, thus the compensation torque for solar pressure is either fully utilized or taken completely away. However, there are times when the solar pressure on the satellite is only partial, as would occur when the satellite is grazing the edge of the earth, such that a full eclipse is never realized. In these cases, the threshold logic circuit 48 may or may not indicate that an eclipse is present depending on whether the threshold value is met and the control torque will either be fully compensated for or not at all even though there is only partial solar pressure on the satellite. It is proposed therefore, to implement another logic circuit which will, instead of looking at a threshold value, have proportionality logic to determine how much current is being used, and thus, attempt to estimate what percentage of the sun is still visible to the satellite and set the compensation torque T I to a value proportional to the full value. The foregoing discussion discloses and describes merely exemplary embodiments of the present invention. One skilled in the art will readily recognize from such discussion, and from the accompanying drawings and claims, that various changes, modifications and variations can be made therein without departing from the scope of the invention as defined in the following claims.

Claims: A method of maintaining an orbiting satellite (10) in a predetermined satellite attitude (Cx ∧; Θ c ) against the disturbance (T d ) of solar pressure on the satellite (10), said method comprising the steps of; measuring the angle (Y; Θ) of the satellite's attitude relative to at least one predetermined axis; applying the measured angle (Y m ; Θ m ) of the at least one predetermined axis to a control system (36); calculating compensating control torques (u cmd ; T c ) in the at least one predetermined axis to correct the attitude of the satellite (10) in the at least one predetermined axis from a predetermined desirable satellite attitude position (Cx ∧; Θ c ); determining whether the satellite (10) is in an eclipse from the sun (22) and applying an eclipse signal indicative of this determination to the control system (36), said control system (36) calculating updated control torques (U cmd ; T c ) in response to the eclipse signal to compensate for the lack of solar pressure on the satellite (10); applying the control torques (U cmd ; T c ) to a satellite actuation device (38); and actuating the satellite (10) in response to the control torques (U cmd ; T c ). The method according to Claim 1 wherein the step of determining whether the satellite (10) is in an eclipse includes the steps of measuring the current through a particular circuit associated with the satellite (10) which has current flow in response to discharge of battery circuits associated with the satellite (10), applying the measured current to a threshold circuit (48) to determine if the measured current exceeded a predetermined threshold level, and applying the eclipse signal to the control system (36) if the measured current exceeds the threshold value. The method according to Claim 1 or claim 2 wherein the step of calculating compensating control torques (U cmd ; T c ) includes the step of applying the measured attitude angle (Θ m ) to a gain system (42) for computing a control torque (T P ) proportional to the measured attitude angle (Θ m ), applying the control torque signal (T P ) to an integration means (44) for integrating the control torque signal (T P ) to develop an integrated signal (T I ), and adding the integrated signal (T I ) to the control torque signal (T P ) to derive a total control torque signal (T c ), said total control torque signal (T c ) be applied to the actuation device (38). The method according to claim 1 or 2 wherein the step of calculating compensating control torques (U cmd ; T C ) includes the step of applying the measured attitude angle (Y m ) to a satellite attitude observer (54) to derive attitude estimates (x ∧) and error signal (y m -cx ∧) between the desired satellite attitude (cx ∧) and the actual satellite attitude (Y m ), applying the error signal (Y m -cx ∧) to a disturbance torque estimator (56) for deriving an estimate (d ∧) of the disturbance torque (T d ) being applied to the satellite (10), applying the disturbance torque signal (d ∧) to the attitude observer (54) and to a control algorithm (58), said control algorithm (58) receiving the control estimate (x ∧) and deriving a control signal (U cmd ) being applied to the actuation device (38). The method according to any of Claims 1, 3 or 4, wherein the step of determining whether the satellite (10) is in an eclipse includes the steps of measuring the current through a circuit associated with the satellite (10) which carries a current indicative of whether batteries in the satellite (10) are discharging and applying the measured current signal to a proportionality logic circuit in order to determine at which rate the batteries are being discharged. The method according to any of Claims 1 - 5, wherein the step of measuring the angle (Y; Θ) to the satellite's attitude includes measuring the satellite's attitude (Y; Θ) by means of a sensor (34) selected from the group consisting of an earth sensor, a star sensor, a sun sensor, a ground beacon sensor and a gyro. A system compensating for solar pressure disturbances (T d ) on an orbiting satellite (10), said system comprising: sensor means (34) for measuring the satellite's attitude (Y; Θ) relative to predetermined satellite axes and providing an attitude measurement signal (Y m ; Θ m ) in the predetermined axes; control means (36) for receiving the attitude measurement signal (Y m ; Θ m ) and calculating satellite attitude compensation control signals (U cmd ; T c ) in the predetermined axes; an eclipse transient compensation device (40) providing an eclipse signal to the control means (36) indicating the satellite (10) is being eclipsed from the sun (22), said control means (36) removing signals (d ∧; T I ) being applied to the actuation means (38) from the control means (36) for compensating for solar pressure on the satellite (10); and actuation means (38) for actuating the satellite (10) in the predetermined axis in response to the compensation control signals (U cmd ; T c ) from the control means (36). The system according to Claim 7 wherein the eclipse transient compensation device (40) is measuring the current flow in a particular circuit associated with the satellite (10), said circuit carrying current indicative of whether a satellite battery is discharging. The system according to Claim 8 wherein the eclipse transient compensation device (40) includes a current measurement device (46) for measuring the current in the particular circuit and a threshold logic circuit (48), said current measurement device (46) applying a current signal to the threshold logic circuit (48) indicative of the level of current being measured by the current measurement device (46) and the threshold logic circuit (48) sending the eclipse signal to the control means (36) if the current signal is above a predetermined threshold value. The system according to Claim 8 wherein the eclipse transient compensation device (40) includes a current measurement device (46) for measuring the current in the particular circuit and a proportionality logic circuit, said current measurement device (46) applying a current signal to the proportionality logic circuit indicative of the level of current being measured by the current measuring device (46) and the proportionality logic circuit sending the eclipse signal to the control means (36) indicative of the extent to which the satellite (10) is being eclipsed from the sun (22). The system according to any of Claims 7 through 10, wherein the sensor (34) is an inertia sensor (34) selected from the group consisting of an earth sensor, a star sensor, a sun sensor, a ground beacon sensor and a gyro, said inertia sensor (34) measuring the angle (y; Θ) of the satellite attitude relative to a predetermined desirable satellite attitude (Cx ∧; Θ c ) along the predetermined axis and outputting a signal (Y m ; Θ m ) representative of these angles (Y; Θ) to the control means (36).

Inventors: BENDER DOUGLAS J, BENDER, DOUGLAS J.
Applicants: HUGHES AIRCRAFT CO, HUGHES AIRCRAFT COMPANY
Application year: 1992
Main classification: B64G1
IPC codes: B64G1, G05D1