Single-axis disturbance compensation system

• US 4,170,904 A
• Filed: 12/12/1977
• Issued: 10/16/1979
• Est. Priority Date: 12/12/1977
• Status: Expired due to Term

• Current Assignee: the united states of america as represented by the secretary of the navy
• Original Assignee: the united states of america as represented by the secretary of the navy
• Inventors: Fischell, Robert E., Sadilek, Albert C., Fountain, Glen H., Mobley, Frederick F.
• Primary Examiner(s): Scott, Samuel
• Assistant Examiner(s): Ferrell, Don E.
• Application Number:
  US05/859,355
• Time in Patent Office: 673 Days
• Field of Search: 244/166, 244/164, 244/170, 244/171, 244/1 R, 244/165, 308/10, 343/DIG. 2, 74/5.5, 74/5.6 R, 74/5.6 A, 74/5.6 D, 74/5.6 E, 74/5.46
• US Class Current: 74/5.60E
• CPC Class Codes: B64G 1/1014 Navigation satellites
• CPC Class Codes: B64G 1/244 Spacecraft control systems
• CPC Class Codes: B64G 1/26 using jets
• CPC Class Codes: B64G 1/285 using momentum wheels
• CPC Class Codes: B64G 1/34 using gravity gradient
• CPC Class Codes: B64G 1/36 using sensors, e.g. sun-sensors, horizon sensors
• CPC Class Codes: G01P 15/08 with conversion into electric or magnetic values
• CPC Class Codes: G01P 15/093 by photoelectric pick-up
• CPC Class Codes: G01P 15/13 by measuring the force required to restore a proofmass subjected to inertial forces to a null position
• CPC Class Codes: Y10T 74/125 by magnetic field
• CPC Class Codes: Y10T 74/1264 Optical
• CPC Class Codes: Y10T 74/1275 Electrical
• CPC Class Codes: Y10T 74/1279 Electrical and magnetic

Abstract

The invention relates generally to an improved system for sensing and compensating for external disturbance forces acting on a satellite while in orbit. A proof mass member is housed within an enclosure and shielded from external, non-gravitational forces. The proof mass is electromagnetically levitated to move in a purely gravitational orbit, along an axis aligned with the satellite'"'"'s velocity vector. The proof mass is subjected to a controlled magnetic biasing field and is caused to have a constant reaction to the resultant biasing force, by means of a thermal control system which maintains constant resistivity of the proof mass, during operation. The position of the proof mass with respect to its axis is detected optically and is utilized to control the firing of spacecraft thrusters. As a result, the satellite is caused to maintain a substantially constant position relative to the proof mass and thereby also is caused to follow a purely gravitational orbit.

Claims

1. In a system for controlling an artificial earth satellite equipped with thruster means to move at a predetermined velocity along a pure gravitational orbit, the improvement comprising,

an enclosure attached to the satellite substantially in alignment with the velocity vector of the orbiting satellite,

an electrically conductive proof mass means disposed within said enclosure and shielded thereby from external non-gravitational forces while said satellite is orbiting,

said enclosure having interior surfaces disposed symmetrically relative to said proof mass means and forming an equipotential surface to shield said proof mass from electrostatic forces,

means for generating a controllable eddy current suspension force on said proof mass means to suspend said proof mass means on a reference axis substantially aligned with the velocity vector of the orbiting satellite,

means for generating a controllable eddy current biasing force on said proof mas means in either direction along said reference axis to position said proof mass means along said axis,

thermal control means for maintaining a constant temperature within said enclosure to keep constant the electrical resistivity of said proof mass means to eddy currents and thereby also the reaction of said proof mass means to said suspension and biasing forces,

means for detecting the position of the proof mass means along said axis, and

thruster control means responsive to said position detecting system for accelerating/decelerating said satellite to maintain a substantially constant position relative to said proof mass means, whereby said satellite also follows a purely gravitational orbit.

2. The satellite control system specified in claim 1 wherein said thermal control means includes a thermistor for monitoring the temperature within said enclosure adjacent said proof mass means, a heater wire, and an energizing control circuit for said heater wire responsive to the temperature monitored by said thermistor.

3. The satellite control system specified in claim 1 wherein

said proof mass means comprises a cylindrical member formed of electrically conductive material and having a central aperture extending through and along the longitudinal axis of the cylinder, and

said means for generating a controllable eddy current suspension force includes a central conductor wire extending through the aperture within said proof mass means along said reference axis, a source of energizing current connected to one end of said central conductor wire, and a plurality of resistive return wires symmetrically disposed relative to said proof mass and connecting the other end of said central conductor wire to said energizing current source.

4. The satellite control system specified in claim 3 wherein said means for detecting the position of the cylindrical proof mass relative to said reference axis comprises an optical system including,

a light source means for transmitting first and second light beams across said reference axis adjacent the opposite ends of said cylindrical proof mass and a third light beam along said reference axis through the central aperture within said cylindrical proof mass, and

light detector means for detecting the amount of each light beam intercepted by said proof mass to determine the longitudinal and transverse positions of said proof mass relative to said reference axis.

5. The satellite control system specified in claim 4 wherein the internal surfaces of said enclosure are optically absorbant to stray radiation from said light source.

6. The system specified in claim 3 wherein said enclosure comprises,

a first hollow cylinder member fabricated of non-magnetic, electrically non-conductive and thermally conductive material, said cylinder having its central longitudinal axis aligned with the velocity vector of the satellite to define a reference axis,

said hollow cylindrical proof mass means being disposed within said first cylinder with its longitudinal axis aligned substantially with said reference axis,

a thermistor disposed to monitor the temperature inside said first cylinder,

a winding mounted on said first cylinder and being energized in response to the temperature monitored by said thermistor to maintain a constant temperature within said first cylinder,

a second cylinder of non-magnetic, electrically non-conductive material adapted to receive and encompass said first cylinder in mating relationship, and

a pair of selectively energizable windings wound around said second cylinder in groups of varying numbers of turns selected to create a linear variation of the square of the magnetic biasing field within said first cylinder along said reference axis effective to exert a controllable constant level biasing force on said proof mass means along said reference axis.

7. The satellite control system specified in claim 6 further including an optical system supported within said first hollow cylinder for detecting the position of said proof mass means relative to said reference axis, said optical system comprising,

a light source means for transmitting first and second light beams across said reference axis adjacent the opposite ends of said cylindrical proof mass and a third light beam along said reference axis through the central aperture within said cylindrical proof mass, and

light detector means for detecting the amount of each light beam intercepted by said proof mass to determine the longitudinal and transverse positions of said proof mass relative to said reference axis, the interior surfaces within said enclosure surrounding said optical system and said proof mass being optical absorbant to stray radiation from said light source and forming an equipotential surface shielding said proof mass from electrostatic forces.

8. The satellite control system specified in claim 1 wherein said means for generating a controllable eddy current biasing force on said proof mass means comprises a pair of energizable windings wound to produce a linear gradient of the square of the magnetic field strength B in opposite directions along said reference axis, and means for selectively energizing said winding pair with a controlled current amplitude.

9. The satellite control system specified in claim 1 further comprising means for retaining said proof mass means in a fixed position until said satellite is placed in its orbital configuration.

10. The satellite control system specified in claim 1 wherein said means for detecting the position of said proof mass means along said reference axis comprises an optical system including a light source means for transmitting light across said reference axis adjacent opposite ends of said proof mass means, and light detector means for detecting the amount of light intercepted by each end of said proof mass means.