## 3 Results
### Experimental Orbit Determination from Interferometric Observables  
The first approach to perform the experimental validation of the system consists of an interferometer with a separation between antennas of 10-15 m with both vertical and horizontal baseline components, as described in Section 2.3. In order to validate the feasibility of the presented technique, an evaluation of the consistency between the estimated observations and the measured ones was carried out. As described in Section 2.2, the process consisted of retrieving the estimated trajectory of the satellite by means of the measured observations and a least-squares estimation technique based on Nelder-Mead optimization, a GEO dynamical model, and a Runge-Kutta-Fehlberg 7(8) integrator. Then, the estimated observations shown in Figure 7 were obtained after computing the range differences between the estimated position of the spacecraft along time and the position of the receiver antennas. The degree of correlation between estimated observations and the measured interferometric observations during the observation window was evaluated. As depicted in Figure 7, one third of the observation window (shaded area) between station-keeping maneuvers was used to retrieve the state vector at the initial time epoch of the observation period (\(t_{0}\)). This state was propagated forward in time by means of the interferometric GEO model and the Runge-Kutta-Fehlberg 7(8) integrator. Since the model does not consider artificial perturbations, the fitting was only valid within an orbital maintenance window of about two weeks. It diverges from the actual motion of the spacecraft when orbital maneuvers are performed. Different orbital solutions as a consequence of the 3D position ambiguity would imply an offset on the average phase value between the estimated and the measured observations in Figure 7. Since the compact baseline presents a distance of ambiguity (60 km) larger than the relative motion of the satellite with respect to the receivers (up to 50 km as shown in Figure 8), the results presented in Figures 7 and 8 correspond to the correct orbital solution. Since the interferometric system provides continuous tracking (one observation per second) of the spacecraft, the orbit determination module is able to make use of updated  
Figure 7: Comparison between the experimental and modeled phase observations. The fitting provides a reference in order to assess whether the interferometric measurements are reliable and consistent. One third of the observation window between orbital maneuvers (shaded area) is used in order to retrieve the orbital state at \(t_{0}\). The model fits the experimental data within an orbital maintenance window. The model does not consider artificial forces and, therefore, diverges from the actual motion of the spacecraft when a station-keeping maneuver is performed (simple orbit determination). The continuous observation of the interferometric signals allows the procedure to use updated data in order to keep track of the satellite even after station-keeping maneuvers (maneuver detection). orbit observations in order to keep track of the spacecraft even after station-keeping maneuvers. Figure 7 shows the result of updating the orbital model with new observations every 24 h (maneuver detection), which yields a much more precise tracking of the satellite. The short-term cycle of the phase, with a daily period, corresponds to the ellipse drawn by the spacecraft due to the inclination and eccentricity of its orbit. In addition, the interferometric observation is capable of tracking the long-term cycle of the spacecraft. With a two-week period, it corresponds to the station-keeping operations, which mitigate the effects of the irregular gravitational potential of the planet. This is a consequence of Earth's nonsphericity and the topographical irregularities of the geoid, creating stable and unstable regions. Any orbiting satellite in a geostationary orbit will tend to deviate to a low geoid.