Hughes Electronics Corporation

Hughes Aircraft Company

The Boeing Company

Jeremiah O Salvatore

Cesar A Ocampo

Stephen C Jennings

Harold A Rosen

Alois Wittmann

Tuan Anh Pham

John R Murphy

Paul D Williams

Christopher D Olsen

Troy D Wallace

A method for operating a satellite () so that a predetermined amount of propellant is left to perform end-of-life maneuvers is provided. A first pressure differential is established between a first propellant tank () and a second propellant tank () by pressurizing the first tank () with a pressurant. Propellant is transferred from the first tank () to the second tank (). Using precise control during transfer the amount of propellant in the tanks may be determined. The second tank is used for orbit maintenance. The first tank is used for end-of-life maneuvers. Propellant may be transferred between the tanks using a latching process so that a predetermined amount of propellant may be transferred from a non-empty tank to a full tank.

Method of operating a satellite for end-of-life maneuvers

A method for operating a satellite ( ) so that a predetermined amount of fuel is left to perform end-of-life maneuvers is provided. A first pressure differential is established between a first propellant tank ( ) and a second propellant tank ( ). Propellant is transferred from the first tank ( ) to the second tank ( ). The second tank is used for orbit maintenance. The first tank is used for end-of-life maneuvers. Propellant may be transferred between the tanks using a latching process so that a predetermined amount of fuel may be transferred from a non-empty tank to a full tank. Another way in which the propellant may be measured is to equalize the amount of propellant in each tank and using one tank to completion so the amount remaining is known.

A satellite system that is coordinatable with a geostationary belt having a plurality of geostationary positions includes an eccentric satellite orbit having a predetermined angle of inclination with respect to the equatorial plane and a non-coincident equatorial crossing node with the geostationary belt. The method of positioning a satellite includes the steps of providing a satellite into a geostationary belt and positioning the satellite in a substantially geosynchronous (24 hr. period) second orbit so that the second orbit is eccentric non-intersecting with the geostationary belt upon the sensing of a predetermined condition indicative of an impending failure on-board the satellite.

Method and apparatus for operating satellites in orbit

The attitude of a spacecraft is controlled during an eclipse without the dependence on auxiliary power by configuring the spacecraft so that it will receive solar power after solar eclipse is passed. The spacecraft is then oriented with respect to a sun line. The spacecraft is then spun so that it maintains a desired angular momentum throughout the eclipse and provides passive stabilization of the attitude. After the eclipse has passed, the spacecraft power is restored and the spacecraft attitude is detected. All spacecraft units are then powered up and thereafter, normal on-station attitude is re-established and normal spacecraft operations are established.

Method and apparatus for using a momentum bias to provide spacecraft attitude stabilization during an eclipse

A method is provided for using at least two lunar flyby maneuvers to transfer a satellite from a quasi-geosynchronous transfer orbit having a high inclination to a final geosynchronous orbit having a low inclination. The invention may be used to take the inclination of a final geosynchronous orbit of a satellite to zero, through the use of a first leading-edge lunar flyby and subsequent successive leading or trailing edge lunar flybys resulting in a geostationary orbit.

Free return lunar flyby transfer method for geosynchronous satellites

A method is provided for using a lunar flyby maneuver to transfer a satellite from a quasi-geosynchronous transfer orbit having a high inclination to a final geosynchronous orbit having a low inclination. The invention may be used to take the inclination of a final geosynchronous orbit of a satellite to zero, resulting in a geostationary orbit, provided that the satellite is launched in March or September.

Spin-stabilized spacecraft are provided which facilitate simple attitude maneuvers and the use of high-gain antennas and efficient solar panels without the need for complex control systems and an excessive number of thrusters. They include a rotor and first and second spacecraft platforms wherein the rotor and the first and second spacecraft platforms are rotatably coupled together and a spin thruster system is arranged to urge and maintain the rotor in a spacecraft-stabilizing spin about the rotor axis. A first rotary-drive mechanism couples the first spacecraft platform to the rotor and a second rotary-drive mechanism couples the second spacecraft platform to the first spacecraft platform. An attitude thruster is spaced from the rotor axis and oriented to generate an axial thrust component in the rotor for synchronous control of the inertial attitude of the rotor axis. A spacecraft of the invention can orbit the earth in an orbital plane and at an orbital rate and be controlled by the steps of (a) spinning the rotor about a rotor axis with a thruster to establish a satellite-stabilizing angular rotor momentum, (b) orienting the rotor axis in an orthogonal relationship with the orbital plane, (c) rotating the first satellite platform substantially at the orbital rate to constantly direct a portion of the first satellite platform at the earth, and (d) counter-rotating the second satellite platform substantially at the orbital rate to constantly direct a portion of the second satellite platform at the sun. The orienting step (b) is simply realized by synchronously generating an axial thrust component in the rotor that is spaced from the rotor axis.

Spin-stabilized spacecraft and methods

A method and system of damping nutation of a spacecraft (20) having a desired spin axis along a first principal inertia axis utilizes a momentum source (28) oriented along a second principal inertia axis perpendicular to the first principal inertia axis. An angular rate of the spacecraft (20) is sensed along an axis transverse to both the first principal inertia axis and the second principal inertia axis. An angular rate signal representative of the angular rate is generated. The angular rate signal is processed to form a control signal representative of a desired torque to drive the momentum source. The desired torque has a first additive component proportional to a derivative of the angular rate to critically damp the nutation under an at least second order model of the spacecraft (20). The momentum source (28) is driven in dependence upon the control signal. Preferably, the desired torque has a second additive component proportional to a momentum stored in the momentum source (28) to inhibit a conversion of the nutation into a tilt of the desired spin axis.

Critical nutation dampling on spinning bodies via momentum wheels or

A method and apparatus for balancing environmental disturbance torques acting upon a satellite (10) by tilting the satellite solar panels (14,16) toward or away from the sun to change the solar pressure acting upon the satellite (10) as well as the gravity gradient torque generated by different panel (14,16) tilt angles. By balancing the disturbance torque acting upon the satellite (10), satellite disturbances are minimized so that the satellite stabilization system can be made smaller and fuel used for satellite attitude control is minimized. Torque balancing is accomplished at the beginning of the mission through a linear actuator (37) at the mounting location of each solar panel (14,16) to the satellite main body (12) and, once set, is passive and does not require continuous active satellite control.

Method and apparatus for satellite torque balancing

A propellant immobilizing system and method in which propellant motion in the propellant storage tanks (22) is reduced or eliminated during velocity change maneuvers to reduce disturbance torques acting upon the spacecraft (10) and improving the spacecraft attitude pointing performance. The thrusters (14, 16, 18 or 20) are fired to produce a small impulse in the direction of the desired Velocity change to begin motion of the propellant within the fuel tank in the opposite direction. After the fuel has moved within the tank to a location in which the propellant center of mass is aligned with the propellant tank (22) center of curvature in the direction of the desired velocity change, thrusters (14, 16, 18 or 20) are again fired to produce a force in the direction of the current velocity of the propellant center of mass to stop the propellant center of mass relative to the propellant tank (22). With the center mass of the propellant now aligned with the propellant tank center of curvature in the direction of the desired velocity change, the thrusters (14, 16, 18 or 20) are again fired to produce the desired velocity change with the propellant center of mass now being trapped in the propellant tank such that additional motion of the propellant is inhibited. After completion of the maneuver, the propellant gradually migrates back to its original position due to the influence of a propellant management device within the propellant tank.

Propellant immobilizing system and method

A spacecraft incorporating momentum sources such as spinning momentum wheels having momentum vectors or components thereof oriented along the desired axis of rotation of the spacecraft. The rotation axis of the spacecraft is nominally any one of the three principal moment of inertia axes of the body. Proper adjustment of the spin rate of the main body relative to the magnitude of the moment sources angular momentum will insure that the spacecraft will spin about the desired axis in a passively stable way even though the rotational axis is not the principal axis of maximum moment of inertia.

Spin stabilization via momentum wheels or similar devices

Spherical tank 30 has its outlet 34 angularly positioned away from the tank center line axis 54 which is parallel to the local gravity vector 26. This angular positioning of outlet 34 inhibits vortexing as liquid is withdrawn from tank 30.

Vortex inhibiting closed spherical outlet structure

A motor active nutation damping system employing a rotor mounted accelerometer, a platform/rotor relative rate sensor, a despin motor and a controller. In the presence of nutation, the accelerometer senses a sinusoidal acceleration whose amplitude is proportional to nutation angle and whose frequency is rotor nutation frequency, i.e., the rate at which the transverse angular momentum vector appears to rotate in rotor fixed coordinates. The controller multiplies the filtered accelerometer signal by a square wave at relative rate generated by the relative rate sensor to obtain a signal consisting of sinusoids at platform nutation frequency and higher order frequencies. Since platform nutation frequency is the rate at which the transverse angular momentum vector appears to rotate in platform fixed coordinates, spin axis torque appropriately phased at this frequency and applied to a dynamically imbalanced platform generates a transverse reaction torque that reduces the transverse angular rate associated with nutation. The controller accomplishes this function by appropriately filtering and amplifying the modulated signal, and then commanding a torque to the despin motor proportional to the resulting signal.

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Redondo Beach, CA, United States of America

Jeremiah O Salvatore