On the satellite attitude determination using simple environmental models and sensor data

Attitude determination represents a fundamental task for most of the spacecrafts. It relies on three basic aspects: 1) sensors selection, 2) relevant environmental conditions estimation, and 3) algorithms that relate the sensor measurements to the expected conditions in the reference frame. Each one has its own impact on the accuracy that the system can achieve. Besides, two factors stand out above the others in terms of accuracy: 1) sensor quality (calibration, range, etc), and 2) precision of the environmental models. The computation of the satellite attitude needs at least two independent measurements (magnetometers, solar sensors...), whit their corresponding simulated measurements in the reference frame. Nevertheless, the number of measurements can be reduced to one if the satellite attitude is constrained. This paper describes a procedure to calculate satellites’ attitude and the main environmental models used (Earth magnetic model, Sun position model, Albedo model), including orbit propagation. This methodology can be extended to measure the performance of a sensor if the satellite attitude can be derived from other measurements and satellite constrains. The methodology is checked with data from the UPMSat-2 mission (launched in September 2020 within the VEGA VV16 mission). This is a 50-kg satellite designed and developed at the Universidad Politécnica de Madrid (UPM).

Re-entry prediction of objects with low-eccentricity orbits based on mean ballistic coefficients

During re-entry objects with low-eccentricity orbits traverse a large portion of the dense atmospheric region almost every orbital revolution. Their perigee decays slowly, but the apogee decays rapidly. Because ballistic coefficients change with altitude, re-entry predictions of objects in low-eccentricity orbits are more difficult than objects in nearly circular orbits. Problems in orbit determination, such as large residuals and non-convergence, arise for this class of objects, especially in the case of sparse observations. In addition, it might be difficult to select suitable initial ballistic coefficient for re-entry prediction. We present a new re-entry prediction method based on mean ballistic coefficients for objects with low-eccentricity orbits. The mean ballistic coefficient reflects the average effect of atmospheric drag during one orbital revolution, and the coefficient is estimated using a semi-numerical method with a step size of one period. The method is tested using Iridium-52 which uses sparse observations as the data source, and ten other objects with low-eccentricity orbits which use TLEs as the data source. We also discuss the performance of the mean ballistic coefficient when used in the evolution of drag characteristics and orbit propagation. The results show that the mean ballistic coefficient is ideal for re-entry prediction and orbit propagation of objects with low-eccentricity orbits.

A Hybrid-Precision Numerical Orbit Integration Technique for Next Generation Gravity Missions

<p>In Next Generation Gravity Missions (NGGM) the Laser Ranging Interferometer (LRI) is applied to measure inter-satellite range rate with nanometer-level precision. Thereby the precision of numerical orbit integration must be higher or at least same as that of LRI and the currently widely-used double-precision orbit integration technique cannot meet the numerical requirements of LRI measurements. Considering quadruple-precision orbit integration arithmetic is time consuming, we propose a hybrid-precision numerical orbit integration technique, in which the double- and quadruple-precision arithmetic is employed in the increment calculation part and orbit propagation part, respectively. Since the round-off errors are not sensitive to the time-demanding increment calculation but to the least time-consuming orbit propagation, the proposed hybrid-precision numerical orbit integration technique is as efficient as the double-precision orbit integration technique, and as precise as the quadruple-precision orbit integration. By using hybrid-precision orbit integration technique, the range rate precision is easily achieved at 10-12m/s in either nominal or Encke form, and furthermore the sub-nanometer-level range precision is obtainable in the Encke form with reference orbit selected as the best-fit one. Therefore, the hybrid-precision orbit integration technique is suggested to be used in the gravity field solutions for NGGM.</p>