Aeronautics and astronautics – Spacecraft – Attitude control
Reexamination Certificate
1997-03-13
2002-05-07
Jordan, Charles T. (Department: 3644)
Aeronautics and astronautics
Spacecraft
Attitude control
Reexamination Certificate
active
06382565
ABSTRACT:
TECHNICAL FIELD
The present invention relates generally to methods and systems for stabilizing the motion of a spacecraft, and more particularly, to methods and systems for stabilizing the rotation of a spacecraft about an intermediate inertia axis.
BACKGROUND OF THE INVENTION
The stability of the rotation of a spacecraft about a desired axis is of concern in man aerospace applications. For example, a transfer orbit spin of a satellite must be stable so that procedures such as attitude determination, thermal control, propellant management, fuel-efficient velocity increment maneuvers, command and telemetry linkage and solar power collection can be accurately performed. When the transfer orbit spin of a satellite is about an intermediate inertia axis, i.e., an axis having a moment of inertia thereabout less than the moment of inertia about a maximum principal axis, and greater than the moment of inertia about a minimum principal axis, the resulting spin is highly unstable. Specifically, a rapidly growing exponential divergence is produced in an uncontrolled intermediate axis spin, as opposed to the slowly-growing divergence which occurs in nutation.
Most geosynchronous communications satellites are of the body-stabilized momentum bias type, and have at least two momentum wheels for providing momentum stabilization on orbit. Such satellites are typically spin-stabilized during transfer orbit, spinning about an axis nearly perpendicular to their momentum wheels. They typically include at least two independent sets of 3-axis gyros to measure body rates to stabilize the satellite during thruster maneuvers during operation.
One solution for obviating the potential for instability is to avoid spinning about an intermediate inertia axis. This can be achieved by imposing constraints in the layout of the satellite in order to produce the desired inertia properties. However, the cost to meet these constraints is excessive as a result of having to produce the desired inertia properties in transfer orbit through deployments to the on-orbit configuration.
Another solution is to employ an active spin axis control system to stabilize the intermediate axis spin. U.S. Pat. No. 4,961,551 to Rosen discloses such a system which uses thrusters under active control with gyro rate sensing. This approach is disadvantageous in that irreplaceable propellant is consumed when using the thrusters, and further, the orbit and momentum of the satellite is disturbed b, use of the thrusters.
U.S. Pat. No. 5,012,992 to Salvatore discloses a system for stabilizing intermediate axis spin which uses two momentum wheels and two gimballed momentum wheel platforms in a “vee wheel” configuration. The momentum wheels and platforms are employed to enhance the spin momentum and make the spin axis appear to have the maximum moment of inertia. A difficulty with this system results from deploying the momentum wheel platforms before the end of deployments, and possibly before the end of LAM firing. Typically, the momentum wheel platforms are delicate mission-critical mechanisms which are not designed to take the resulting deployment/LAM loads. Also, since both of the momentum wheels and platforms are utilized, the resulting system is not single fault tolerant.
SUMMARY OF THE INVENTION
It is an object of the present invention to stabilize an intermediate inertia axis spin of a spacecraft without employing thrusters.
Another object of the present invention is to stabilize an intermediate inertia axis spin without using gimballed momentum wheel platforms.
A further object of the present invention is to provide a system for stabilizing an intermediate inertia axis spin which is single fault tolerant.
In carrying out the above objects, the present invention provides a method of stabilizing the spin of a satellite about an intermediate inertia axis. A first component of angular velocity of the satellite about a first axis transverse to the intermediate inertia axis is sensed. A second component of angular velocity of the satellite about a second axis transverse to the intermediate inertia axis is also sensed. A single degree of freedom momentum storage device, which has a fixed transverse orientation with respect to the intermediate inertia axis, is controlled in dependence upon the first component and the second component of angular velocity.
Further in carrying out the above objects, the present invention provides a method of stabilizing the spin of a satellite about an intermediate inertia axis. A first component of angular velocity of the satellite about a first axis transverse to the intermediate inertia axis, and a second component of angular velocity of the satellite about a second axis transverse to the intermediate axis are sensed. The second axis is transverse with respect to the first axis. The angular velocity of a single degree of freedom momentum wheel having a fixed transverse orientation with respect to the intermediate inertia axis is sensed. A commanded angular velocity for the momentum wheel is computed in dependence upon the first component and the second component of angular velocity. The commanded angular velocity is proportionate to a transverse component of the angular velocity of the satellite, wherein the transverse component is in a direction based upon an eigenvector associated with an unstable eigenvalue of a linear dynamic model of the satellite. A control signal representative of a control torque to be applied to the momentum wheel is formed. The control torque is proportional to a difference between the angular velocity of the momentum wheel and the commanded angular velocity. The control signal is applied to the momentum wheel.
Still further in carrying out the above objects, the present invention provides a system for stabilizing the spin of a satellite about an intermediate inertia axis. A pair of gyros sense a first component of angular velocity about a first axis of the satellite, and a second component of angular velocity about a second axis of the satellite, wherein the first axis and the second axis are each transverse to the intermediate inertia axis. A processor forms a control signal in dependence upon the first component and the second component of angular velocity. The control signal is applied to a single degree of freedom momentum storage device having a fixed transverse orientation with respect to the intermediate inertia axis.
These and other features, aspects, and advantages of the present invention will become better understood with regard to the following description, appended claims, and accompanying drawings.
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J.J. Adams, “Study of an Active Control System for a Spinning Body”, NASA TN D-905, Technical Note, Langley Research Center, VA, NASA Jun. 1961, pp. 1-31.
J.J. Adams, “Simulator Study of an Active Control System for a Spinning Body”, NASA TN D-1515, Langley Research Center, Dec. 1962, pp. 1-22.
M. Loebel; “Several Linear Stabilization and Reorientation Control System Configurations for a Rotating, Manned Orbital Space Station”; Academic Press, Guidance and Control-II; AIAA Guidance and Control Conference, Cambridge, Mass, Aug. 12-14, 1963; pp. 313-337.
N.H. Beachley and J.J. Uicker, Jr., “Wobble Spin Technique for Spacecraft Inversion and Earth Photography”, Journal of Spacecraft, vol. 6, No. 3,; Mar. 1969, pp. 245-248.
T.R. Kand & M.P. Scher, “A Method of Active Attitude Control Based on Energy Considerations”, J. Spacecraft
Dinh Tien
Gudmestad Terje
Hughes Electronics Corporation
Jordan Charles T.
LandOfFree
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