Abstract: A three-axis stabilized spacecraft is subject to a velocity change in a desired direction by a thruster. Unavoidable alignment errors cause a body torque, which tends to slew the body away from the attitude which orients the thrust axis in the direction. The attitude control system eventually corrects the attitude, but the thrust during the attitude transient results in an error in the direction of the velocity change. This error in the direction accumulates during the attitude transient. When the attitude transient passes, no further pointing error occurs. A control system, operating without an accelerometer, determines the total error accumulated during the attitude transient, and processes the error signal to generate a supplemental torque demand signal, which is added to the torque demand signal produced by the attitude control system, to cause an oppositely-directed attitude transient to thereby cancel the original velocity change direction error.

Abstract: A spacecraft (10) carries a solar panel (17) which rotates to follow the sun, and also carries various thrusters (20). Thruster plume impingement on the solar panel affects the torque applied to the spacecraft body (12) in a manner which depends upon solar panel angle. The errors in the thrust during stationkeeping tend to perturb attitude, especially early in the maneuver, because of the delay inherent in the attitude control loop. A torque bias is summed with the residual torque demand signal to correct for the errors in torque. The torque bias signal is generated by a Fourier model of the torques, updated by an adaptive tuning filter, so that successive stationkeeping maneuvers progressively adapt the amplitude and phase of the Fourier coefficients in a manner which tends to minimize the residual torque demand and attitude error. Thus, the torque bias signal automatically approaches the correct value.

Abstract: A spacecraft (10) has an attitude sensor (20) mounted on its body (12). The sensor must be maintained near a temperature setpoint. Each sensor (20) produces its own temperature-indicative signal. Each attitude sensor is coupled to a thermally conductive instrument platform (18). A standoff (21) supports the platform (18) away from a baseplate (16) and the spacecraft body (12). The standoff (21) includes a thermally conductive portion (22) adjacent the platform (18), and a nonconductive portion (24) remote from the platform. An electric heater (26) is connected to the thermally conductive portion (22) of the standoff (21). A temperature sensor (28) thermally coupled to the platform (18) generates platform temperature signals. A filter (212) high-pass filters either the platform temperature signals or the attitude sensor temperature signals, to form filtered signals. A combining circuit (218) combines the filtered signals with the other temperature signals, to make composite temperature signals.

Abstract: A spacecraft includes a three-axis attitude control system. When velocity change thrusters are fired, their plumes impinge on a solar array, at angles which vary with the solar array position. This causes disturbance torques which vary with the solar array position. Disturbance torque information signals or torque bias signals which depend upon the solar array angle are summed with the torque demand signals which control the attitude control system during firing of the velocity change thrusters, to modify the attitude correction torques. The bias torque signals are generated by a Fourier processor based upon stored Fourier coefficients together with signals from a solar array angular position sensor.

Abstract: A three-axis stabilized spacecraft includes a plurality of primary attitude control thrusters, the torque vectors of which lie in, or parallel to a primary plane. It also includes at least two more secondary attitude control thrusters, the torque vectors of which lie in a secondary plane which is not parallel to the primary plane. The control system produces attitude error signals, which are processed with a PID characteristic to produce impulse demand signals, all in known fashion. The impulse demand signals are transformed into an auxiliary coordinate system, in which two of the three auxiliary axes lie in the primary plane, and the third is orthogonal thereto. One of the secondary thrusters is selected, which has, along the third auxiliary axis, the largest torque magnitude and the same sign as the transformed impulse demand.