An ultra low weight/low cost three axis attitude readout system for nano-satellites

This paper describes a +/- 1.5 degrees accuracy, low weight/power/cost, three axis attitude readout sensor system designed to local-vertical-stabilize a nano-satellite in a low earth orbit. This is accomplished with one or two, +/-0.5 degrees sun sensors, and a two-degree-of-freedom gyro. These sun sensors allow a sun view for a portion of the sun-illuminated portion of a wide variety of orbit inclinations, thus updating the gyro drift. For sun synchronous orbits, or those with minimum inclination, only one sun sensor is required. Each sun sensor weighs less than 34 grams and uses milliwatts of power. The sun sensor generates a two-axis readout of the sun's angular position anywhere within its 100 degrees diameter field of view. This is accomplished by using the spacecraft computer to interpolate the sensor's real time analog outputs within a precision digital lookup table of the two axis sun sensor's analog outputs, generated by a factory calibration. Analysis and test of prototypes of this sun sensor has shown that it provides a sun position readout accuracy of better than +/-0.5 degrees. This attitude readout capability is maintained during the dark portion of the orbit by a MINITACT (TM) two degree of freedom gyro. The MINITACT (TM) gyro is currently used in quantity in tactical missiles, and has extremely rugged environmental survival and operation capabilities. It incorporates sealed air bearing rotors, thus achieving unlimited life, and has drift rates less than 0.5 degrees /hour. These gyros and their electronics weigh less than 150 gram each, and consume only 6 watts of power. Therefore, the entire attitude readout system weighs less than 215 grams, and consumes 6 watts of power. The gyro maintains the +/-0.5 degrees pitch and roll attitude readout accuracy during the dark portion of the orbit, and provides the yaw readout based upon its pitch and roll rates. The gyro drift is periodically corrected once per orbit during the sun illuminated portion by one or both of the sun sensors to maintain the +/-1 degrees attitude readout capability indefinitely. The orbit ephemeris is stored in the spacecraft computer. The attitude determination system is based on an extended Kalman Filter. The plant model used by the Kalman Filter includes a model of the attitude dynamics and a gyro noise model. The sun sensor measurements are incorporated as measurements into the filter. The filter estimates three axis attitude, gyro biases and external disturbances. This attitude determination algorithm is described in the paper. The paper also includes an accuracy analysis of the sun sensor, including earth albedo error minimization techniques, if applicable.