Abstract:
An onboard attitude control system is constructed to utilize a four reaction wheel system having a reference axis, wherein at least three of the reaction wheel spin axes are oriented obliquely to the reference axis. Current attitude is estimated based on uploaded orbital data, onboard sensed earth and sun position data, and attitude data sensed by a three axes gyroscope system. Current attitude is compared to mission attitude to calculate an error which is transformed to a trihedral axes adjustment command to actuate the reaction wheel system.

Description:
BACKGROUND OF THE INVENTION 
     The system of this application is designed to control the attitude of a satellite. Satellites must be maintained in a predetermined orbit and attitude in order to accomplish the assigned mission which can be surveillance, photography, detection and many others. The orbit and attitude of the satellite must be periodically adjusted to compensate for disturbances which occur in space or for the purpose of changing the mission. 
     In general, spacecraft attitude is adjusted by activating actuators, such as, momentum wheels, magnetic torguers, or thrusters in response to an attitude correction signal. The attitude error may be sensed by reference to sensors monitoring the position of the sun, stars and earth relative to the satellite or by onboard inertial sensors such as gyroscopes. The attitude is adjusted to its mission orientation in which the system is pointed at its predetermined target and is maintained in this orientation during orbital flight. During flight the satellite is subject to motions induced by external forces, on board mechanisms or other sources and the attitude control system must continuously monitor and adjust attitude. Attitude control is therefore of primary importance in order to point the satellite to accomplish its mission and to maintain that position with the required accuracy. 
     It is a purpose of this invention to use fuel efficient reaction wheels to achieve attitude corrections wherever possible. This is accomplished while obtaining a tighter pointing capability with increased spacecraft autonomy. It is a purpose of this invention to utilize four reaction wheels preferably arranged in a trihedral configuration to provide better performance and enhanced redundancy. 
     SUMMARY OF THE INVENTION 
     The control system of this invention employs a configuration of four reaction wheels preferably arranged in a trihedral relation as the primary attitude adjustment mechanism. The primary attitude sensors consist of a three axis gyroscope system. The control module includes stored orbital and related sun ephemeris data and appropriate estimating algorithms. The attitude is estimated with reference to the output of the three axis gyroscope system. Compensation for errors relating to gyroscope drift are provided by reference to data from on board earth and star sensors. Utilizing this data, the control module provides an estimate of the actual attitude of the satellite. The estimated attitude is compared to the desired attitude to obtain an attitude adjustment in terms of the three axis reference system. Each of the components of the adjustment is transformed to obtain the four wheel torque rates required to accomplish the adjustment. In order to further refine the estimated adjustment, the predicted three axis torque adjustments are fed back to the control module. 
    
    
     DESCRIPTION OF THE DRAWING 
     The invention of this application is described in more detail below with reference to the Drawing in which: 
     FIG. 1 is a schematic illustration of a satellite system using this invention. 
     FIG. 2 is a block diagram of the system of this invention; 
     FIG. 3 is a schematic illustration of the trihedral reaction wheel configuration of this invention; and 
     FIG. 4 is a schematic illustration of an alternative reaction wheel configuration used in this invention. 
    
    
     DESCRIPTION OF THE PREFERRED EMBODIMENT 
     The basic components of the satellite  1  are shown in FIG.  1  and include mission sensors  2 , ground control computer  3 , attitude adjustment actuators  4 , and onboard computer  5 . Ground control computer  3  is in microwave communication with the satellite computer  5  and transmits the attitude data required to fulfill the mission for storage in computer  5 . The mission sensors  2  may include cameras, telescopes, communications antennae and other similar devices. The attitude adjustment actuators  4  are shown schematically as thrusters, but may also include other types of actuators, such as thrusters, momentum wheels, or magnetic torquers. In particular a system of reaction wheels  18 , as shown in FIG. 3, are used to supplement the thruster system. The thrusters  4  may be used for orbit transfer maneuvers, while the reaction wheel system  18  is used for smaller station keeping adjustments. 
     The satellite  1  is oriented in space by reference to three axes orthogonal coordinates. As shown in FIG. 1, the coordinate system includes an x axis which is generally tangent to the orbit path and referred to as the roll axis, a z axis which is generally pointed at the center of the earth and referred to as the yaw axis, and a y axis which is perpendicular to the other axes and referred to as the pitch axis. Pitch, yaw, and roll refer to rotational movement of the satellite about the particular axis. 
     In order to maintain the desired mission attitude a continuous monitoring of actual attitude needs to be accomplished. For this purpose an onboard attitude control module  7  is constructed as part of the satellite control computer  5 . A block diagram of the attitude control module  7  is shown in FIG.  2 . The estimator module  8  contains modeling software which is capable of estimating the actual attitude of the satellite  1  from data sensed on board. Ephemeris and orbital data is up loaded and stored in the attitude control module  7  to allow the modeling software to take into consideration repetitive error causing disturbances. The primary source of sensed attitude data is a three axis gyroscope assembly  9 . The data sensed by the gyroscopes are fed to the estimator module  8  and used to obtain an updated attitude for the satellite  1 . To allow the estimator module  8  to compensating for gyro drift, the position of the satellite  1  with respect to the earth and sun are sensed by earth sensors  14  and sun sensors  15  on the satellite  1 . Data from these sensors are sent to the estimator module  8  and factored into the modeling calculations. The estimator module uses least-square estimation techniques to combine the gyro data with the earth sensor data and sun sensor data to estimate both spacecraft attitude and gyro drift. 
     The modeling software may be any of the available algorithms designed to calculate attitude from available data. 
     The desired mission attitude is periodically up loaded from ground control computer  3  and stored in the attitude control module  7 . The estimated actual attitude is compared to the mission attitude by the adjustment module  16  and an error calculation is obtained. This error calculation is converted to an attitude adjustment with components referencing the standard coordinate system. These data is converted by algorithms in the torque transformation module  17  to a four axis adjustment for actuating the four wheel actuator system  18 . The latter transformation is accomplished as described in U.S. Pat. No. 5,826,829, which issued on Oct. 27, 1998, the contents of which are incorporated herein by reference. 
     The trihedral momentum bias (TMB) wheel configuration of the invention uses four wheels of which any three can be used to provide the momentum bias and active nadir attitude three axis pointing. The four wheels are comprised of one momentum wheel and three reaction wheels. The three reaction wheels (typically smaller than the momentum wheel) are in a trihedral configuration which can provide the backup momentum bias should the momentum wheel fail. Full three-axis control would also be maintained if any one of the reaction wheels should fail. The wheel system can be operated in any of five modes: one using all four wheels and four modes each of which turn off one of the four wheels. It is up to the user which of five available wheel combinations will be used for nominal operation. If the three reaction wheels are used for nominal operations. and the reaction wheels are sized properly, it is possible to achieve three-axis active attitude control without any wheels being required to spin through zero rpm. 
     The trihedral wheel system  18  includes a relatively large momentum wheel  10  mounted on the satellite, which wheel is rotatable about a spin axis (not shown) for maintaining gyroscopic stiffness of the spacecraft in space about a first axis. 
     The wheel system  18  also includes a plurality of relatively smaller reaction wheels  11 ,  12 , and  13  which, like the momentum wheel  10 , are mounted on the spacecraft and rotatable on spin axes  111 ,  112 , and  113 , respectively, in a fixed, trihedral, configuration. Any two of the three reaction wheels  11 - 13 , together with the momentum wheel  10 , provide full three-axis control of the spacecraft in a predetermined attitude. The reaction wheels are flywheels with a vehicle-fixed axis designed to operate through zero wheel speed. In the event of a failure of the momentum wheel  10 , the reaction wheels  11 ,  12 , and  13  can be used to provide angular momentum sufficient to maintain the gyroscopic stiffness lost by the failure of the momentum wheel, while maintaining full three-axis control of the spacecraft in a predetermined attitude. Further, in the event of a failure of any of the wheels, the combined angular momentum of the remaining wheels is effective to maintain gyroscopic stiffness about the first axis while also maintaining full three-axis control of the spacecraft in a predetermined attitude. In short, the momentum wheel  10  and the reaction wheels  11 ,  12 ,  13  are all rotatable about relatively fixed spin axes  111 ,  112 , and  113  in a configuration for together maintaining gyroscopic stiffness and for maintaining three-axis control of the spacecraft. The details of this configuration are described in the above cited patent which is incorporated herein. 
     An alternative embodiment to the trihedral reaction wheel configuration, described above, is shown in FIG.  4 . In this four wheel reaction wheel system  19 , reaction wheels  10 - 13  are arranged with their spin axis  110  through  113  oblique to the y axis. As shown, the angle of each of the spin axes  110 - 113  is approximately 35°, but this could be virtually any angle depending on the amount of momentum bias needed for gyroscopic stiffness and the amount of momentum storage needed. Two of the spin axes are positioned in the yz plane and the other two spin axes are in the xy plane. Similarly to the trihedral configuration, the double V type of configuration also provides three axis control and three axis momentum storage by using any three of the four reaction wheels, while maintaining gyroscopic stiffness.