# A gyrocompassing spacecraft navigator

- Paper ID
1968-AT67

- author
- company
Autonetics, a division of North American-Rockwell Corp.

- country
U.S.A.

- year
1968

- abstract
A new onboard orbit determination method , in which conic orbit parameters are determined from measurements of the orbit angular velocity, is presented. Equations are developed which uniquely express the orbit parameters (such as radius, eccentricity, true anomaly, etc.) in terms of three angular motion variables, the orbit angular velocity and its first two derivatives. Statistical estimation methods are applied to formulate two navigation techniques based upon these equations, The first technique is explicit, that is, free of dependence upon a preplanned nominal trajectory . A statistical filter fits a sequence of noisy angular velocity measurements from the instrumentation system to a dynamic model of the orbit angular velocity, producing estimates of the angular motion variables at a point in the orbit. These estimates are then substituted into the explicit equations to produce orbit parameter estimates directly. The second technique is implicit, and a statistical filter estimates the deviations of the orbit parameters from expected nominal values. A gyrocompass system is proposed to instrument high quality measurements of the orbit angular velocity. Gyrocompass operation in orbit is shown to be directly analogous to operation at a fixed location on the earth equator, except that horizon scanners serve in place of accelerometers to track the vertical and the orbit angular velocity is generally time-varying, A proposed configuration for the system is described, Design features include adaptive system gain computations to compensate for the time-varying angular velocity and a gyro drift rate bias calibration technique to remove angular velocity measurement bias. The orbit determination method is intended especially for spacecraft flying in the near vicinity of a planet, and the performance of the method has been simulated for both captured and flyby orbits about the earth, Data from the simulation studies show that effective navigation can take place while a spacecraft travels a relatively short orbit segment, ranging from about 20 to about 50 degrees of central angle. The permissable measurement noise level is shown to be similar to that of high quality inertial terrestrial navigators , so that the use of inertial instruments is within the realm of possibility The data also support the possibility that the explicit and implicit techniques can be used together to achieve high navigation accuracy with complete freedom from preplanned nominal flight conditions, and the necessary navigation computations are shown to be relatively simple.