Modeling and Design of an Orbital Simulator and of Guidance Algorithms for VLEO maneuvers

With the ever growing interest in space operations more and more studies are considering Very Low Earth Orbits (VLEO) for specific mission; due to the lower altitude than traditional ones these types of orbits have different benefits and drawbacks. The advantages of this orbits are reduced communication latency, improved link budgets and enhanced ground resolution for Earth Observation missions. Since the launcher has to reach lower altitudes the launch costs are reduced and, due to new laws for end-of-life disposals, the thicker atmosphere helps in de-orbiting spacecrafts naturally in reduced time. The lower altitude also has its challenges, i.e. the gravitational anomalies are more severe, the aerodynamic drag is orders of magnitude bigger than at higher orbits and the interactions with the magnetosphere are more pronounced. The objective of this thesis is to enable feasibility analysis of missions in VLEO, a 6 degree-of-freedom simulator is designed in MATLAB/Simulink and incorporates the models of the primary disturbances affecting satellite dynamics in an Earth-Centered Inertial (ECI) frame. The orbit considered in this study is a near-circular Sun-synchronous orbit at an altitude of approximately 300 km. The sources for disturbances considered in this work are: aerodynamic drag, J2 Perturbations and Solar Radiation Pressure, considering a conical shape umbra from Earth. The aerodynamic drag is the greater disturbance in VLEO; the chosen model is Harris-Priester, in which the fidelity is increased with no increase in computational time. The second disturbance (J2 perturbation) is due to the oblateness of the Earth are the highest impacting disturbance on the position dynamics. After the modeling of the primary disturbances the following step in the thesis was the design of an orbital guidance and control algorithm, to prove the concept of the simulator; firstly the guidance system was based on an Artificial Potential Field (APF), where the desired position is determined by propagating the satellite’s initial conditions along an unperturbed Keplerian orbit. This reference trajectory is used as the target state for the APF, which computes a desired velocity, effectively steering the spacecraft toward the intended orbital path. To execute this guidance command, a Sliding Mode Control (SMC) algorithm was employed, enabling robust handling of system non-linearities and external disturbances when operating with discrete cold gas thrusters. A Model Predictive Control (MPC) strategy was then integrated for comparative analysis, leveraging its predictive capabilities to optimize control actions over a finite horizon while accounting for system constraints and actuator limitations. In the absence of active attitude and orbit control, and accounting for the perturbative effects described above, the spacecraft would experience rapid orbital decay, by implementing the proposed control strategy in conjunction with actuators models orbital maintenance is achievable, albeit at a significant propellant cost.