Classifications

• • B—PERFORMING OPERATIONS; TRANSPORTING
  • B64—AIRCRAFT; AVIATION; COSMONAUTICS
  • B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
  • B64G1/00—Cosmonautic vehicles
  • B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
  • B64G1/24—Guiding or controlling apparatus, e.g. for attitude control
  • B64G1/28—Guiding or controlling apparatus, e.g. for attitude control using inertia or gyro effect
  • B64G1/286—Guiding or controlling apparatus, e.g. for attitude control using inertia or gyro effect using control momentum gyroscopes (CMGs)
• • B—PERFORMING OPERATIONS; TRANSPORTING
  • B64—AIRCRAFT; AVIATION; COSMONAUTICS
  • B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
  • B64G1/00—Cosmonautic vehicles
  • B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
  • B64G1/24—Guiding or controlling apparatus, e.g. for attitude control
  • B64G1/244—Spacecraft control systems
• • B—PERFORMING OPERATIONS; TRANSPORTING
  • B64—AIRCRAFT; AVIATION; COSMONAUTICS
  • B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
  • B64G1/00—Cosmonautic vehicles
  • B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
  • B64G1/24—Guiding or controlling apparatus, e.g. for attitude control
  • B64G1/244—Spacecraft control systems
  • B64G1/245—Attitude control algorithms for spacecraft attitude control
• • B—PERFORMING OPERATIONS; TRANSPORTING
  • B64—AIRCRAFT; AVIATION; COSMONAUTICS
  • B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
  • B64G1/00—Cosmonautic vehicles
  • B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
  • B64G1/24—Guiding or controlling apparatus, e.g. for attitude control
  • B64G1/36—Guiding or controlling apparatus, e.g. for attitude control using sensors, e.g. sun-sensors, horizon sensors

Definitions

• This invention relates to satellites and robotic systems, for example controlling the orientation of a satellite using a plurality of control moment gyros (CMG).
• CMG control moment gyros
• attitude of an agile spacecraft or satellite is often maintained and adjusted with a control moment gyro array because those devices provide high torque and torque amplification.
• a typical CMG is a rotating mass suspended on a gimbal with an actuator to rotate it on the gimbal axis, producing torque and accumulating angular momentum.
• Angular momentum is the integral of torque over time.
• An array of «>3 CMGs is often used, allowing attitude control with some redundancy.
• Each CMG has an angular momentum (h) constrained essentially to a plane, the angular momentum vector of the gyro is nearly orthogonal to the gimbal axis.
• the error in orthogonality is small enough that it does not affect the operation of the CMG, the array of CMGs, or the attitude control of the satellite.
• the wheel speed of the CMG is essentially constant in most applications, but does not have to be for this invention to work.
• attitude control calculates the desired attitude rates for the satellite d , being the three axis attitude rates.
• the gimbal angle ( ⁇ ) rates for the CMG array are calculated using the pseudo inverse control law,
• An object of the present invention is to significantly increase the speed in reorienting a satellite between two objects by utilizing more of the available angular momentum from the CMGs.
• a disturbance is introduced into the torque command to cause the CMG array to avoid the singularity.
• the value of the determinant (AA T ) is monitored continuously during CMG operation.
• the required torque command is altered so as to allow the system to escape the singularity.
• the torque can be altered by adding a small fixed amount of torque in one or more of the axes, or by
• a deviation in the torque command is perceived as a disturbance to the spacecraft Inertial Measurement Unit (IMU) which subsequently issues updated torque commands to correct for the disturbance.
• IMU spacecraft Inertial Measurement Unit
• Fig 1. is a functional block diagram showing a control embodying the present invention to rotate a satellite in response to commanded rotation signal q c .
• Fig 2 is a block diagram showing a satellite with CMGs that are rotated to change the satellite's attitude in response to individually produce angular rate signals.
• Fig. 3 illustrates two possible paths for reorienting between two objects.
• Fig. 1 shows function blocks that may be implemented through hardware or software, preferably the latter in a computer based satellite control containing one or more signal processors programmed to produce output signals to control CMGs on the satellite as explained hereafter. Fundamentally the process is shown for a single signal path between two points, but it should be understood that single lines represent vector data which is 3 dimensional for the satellite attitude, attitude rate and torques, and n dimensional for the signals related to the n CMGs.
• a typical fully closed loop control follows an eigen axis path "old" by controlling the CMG's based on the actual attitude (determined from the attitude determination system ADS in Fig. 3) and desired path attitude.
• a desired spacecraft attitude 10 is compared with an actual attitude 12, in a summation junction 14.
• the attitude error 16 from the summing junction 14 is used by the spacecraft attitude controller 18 to produce a desired spacecraft acceleration 20.
• the desired acceleration 20 is multiplied by the spacecraft inertia matrix 22 to produce a torque command 24.
• This torque command 24 is applied to a singularity avoidance process 32 along with the values of the Jacobian matrix A 40 where it is decided if a singularity is being encountered and, if a singularity is eminent, a small torque is calculated 26 that enables CMG array singularity escape which is then sent to a summing junction 28.
• the modified torque command at 34 is sent to the linear transform in box 36 where a modified Moore-Penrose Pseudo-Inverse is utilized to calculate the CMG gimbal angle rate commands 42.
• the gimbal angle rate commands 42 operate the CMG array 48 which produce torque 50 which acts on the spacecraft 52 to change its rate of rotation 54.
• the gimbal angles from the array 46 are used to calculate the values of the Jacobian A 40 at step 44 which is used by boxes 32 and 36.
• the actual rate of spacecraft rotation 54 is measured by sensors 56 which is used to determine the actual spacecraft attitude 12.
• the singularity avoidance process 32 in general, a) determines if there is a potential that a singularity will be encountered (i.e., det (AA T ) ⁇ tolerance), b) if a singularity is likely to be encountered, a small torque that will cause singularity escape is calculated and added to the existing torque command 24 c) otherwise, the torque command is left alone.
• the additional torque command 26 necessary to escape the singularity is small compared to the command 24 and the vector representing this additional torque 26 cannot be colinear with the existing torque command 24. This process enables the CMG array to be utilized to the maximum available momentum envelope for the given array configuration.

Landscapes

• Engineering & Computer Science (AREA)
• Remote Sensing (AREA)
• Chemical & Material Sciences (AREA)
• Combustion & Propulsion (AREA)
• Radar, Positioning & Navigation (AREA)
• Aviation & Aerospace Engineering (AREA)
• Automation & Control Theory (AREA)
• Control Of Position, Course, Altitude, Or Attitude Of Moving Bodies (AREA)

Priority Applications (3)

Applications Claiming Priority (2)

Publications (2)

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Family Applications (1)

Country Status (5)

Cited By (2)

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• 1998
  • 1998-03-16 US US09/039,640 patent/US6047927A/en not_active Expired - Lifetime
• 1999
  • 1999-03-16 DE DE69925719T patent/DE69925719T2/de not_active Expired - Lifetime
  • 1999-03-16 EP EP99935274A patent/EP1064196B1/en not_active Expired - Lifetime
  • 1999-03-16 WO PCT/US1999/005598 patent/WO1999050144A2/en not_active Ceased
  • 1999-03-16 JP JP2000541068A patent/JP4249902B2/ja not_active Expired - Fee Related

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