Development of the navigational system for a multi-spectral imaging nano-satellite

The navigational system of the satellite is required for determination of the position and velocity of a satellite. This is required for obtaining outputs from ephemeris models such as IGRF, WMM and Earth reflectivity for albedo. Given the high pointing requirements for an imaging payload, there exists a need for highly accurate position and velocity information as it affords greater room for inaccuracies to the entire attitude determination, estimation and control system. The navigation system comprises 2 major segments - orbit determination and orbit propagation. The orbit determination segment is responsible for calculating the current position and velocity vectors which is to be carried out using GNSS/GPS receivers. The data from the GNSS receivers provides measurements with errors within acceptable limits given the constraints by our payload. The data obtained from orbit determination is used for initialisation in orbit prop-agation. The in-house orbit propagator has the responsibility of providing an estimated satellite position and velocity at any future point in time along with the reference frame vectors such as the magnetic field vector and Sun vector, which would be utilised in the attitude estimation assembly and power production. This information would in turn be useful for switching between the satellite's various modes of operation. The orbit propagator will be implemented in C due to faster run time and ease of implementation on microcontrollers. The accuracy of this system is verified against data from proven softwares such as NASA's General Mission Analysis Tool (GMAT). We will also be calculating the magnetic field vector field and sun vector using our current position and power generated using the sun vector and the solar panel specifications. There also arises a need to create a system of repeated propa-gation, given the increase in error as the satellite state is propa-gated further into the future. From our pointing requirements (obtained from payload) and the pointing error, we can calculate the tolerable error in position. Using this tolerable error and the modes of operations, we will then develop an algorithm which will optimise the usage of GPS receivers, as they consume significant energy. The exact specifications for such a system are also being investigated. The Orbit propagator incorporates J2 perturbations, solar ra-diation pressure and atmospheric drag. Although various other perturbing forces act on the satellite as well such as gravitational effects from other celestial bodies, higher order perturbations from the Earth's gravitational field among others, the effect of these forces is considered negligible compared to the forces which are included. Various numerical schemes for propagation of satellite state were also investigated in order to find an optimal balance between accuracy and computational requirements.