Bogdan Udrea

Collision avoidance in satellite clusters

A collision avoidance methodology for satellite clusters is developed and simulated. A set membership algorithm is used to monitor and predict guaranteed position uncertainty ellipsoids over time. Distances between ellipsoids can be calculated on-line in order to evaluate when and how fast a collision may occur. The algorithm assumes guaranteed knowledge of an initial state (relative position and velocity) error, sensor error, and forcing function. The forcing function is variable depending on orbital parameters and faults. A probability of collision is evaluated using the resulting ellipsoid information. Results show that satellite clusters with a close proximity will require a back-up system that is accurate and gives measurement updates at least every few minutes.

A cooperative multi-satellite mission for controlled active debris removal from low Earth orbit

This paper presents the concept of operations and preliminary design of a multi-satellite mission for the active removal of large pieces of debris from low Earth orbit. The mission consists of a mothership minisatellite, that carries six nanosatellites. The mothership acquires a relative orbit of a few kilometers with respect to the piece of orbital debris of interest and determines the attitude state of the debris and good docking locations for the nanosatellites. The nanosatellites deploy sequentially and dock with the piece of debris. Once all the nanosatellites are docked with the debris they cooperatively perform its structural analysis to determine safe maneuvering profiles for its detumble and deorbit. The mothership then docks with the piece of debris and applies maneuvers to deorbit it. Systems engineering budgets have been determined for the mass and propellant and a preliminary mission cost has been estimated. They are presented together with the functional architectures of each spacecraft and the results of mission design obtained with the Systems Tool Kit.

OPTIMAL TRAJECTORY PLANNING OF FORMATION FLYING SPACECRAFT

Abstract This paper introduces a closed-loop Guidance and Control optimal algorithm that balances the propellant consumption and the need for collision avoidance among formation flying spacecraft. This model-based algorithm is purely algebraic and computes the spacecraft trajectories from the knowledge of the formation linearized relative dynamics equations and the formation full state. Using Pontryagins maximum principle, Guidance generates the control inputs required to obtain the optimal trajectory from the current state until the target state, and does so for each of a set of regularly spaced time instants. After each recomputation, Control applies the optimal inputs until the next regularly spaced time instant, when Guidance updates the optimal trajectory again. Simulation results for the algorithm applied to a 3-spacecraft formation in GTO are presented.

An approximate Riemann solver for MHD computations on parallel architectures

An implicit algorithm to model time-dependent magnetohydrodynamics (MHD) in three dimensions is described. The ideal MHD equations are hyperbolic, but the addition of resistivity and viscosity changes the mathematical form to mixed hyperbolic-parabolic. The algorithm is a finite volume implementation of a Roe-type approximate Riemann solver combined with a symmetric Gauss-Seidel method. The algorithm is implemented on parallel architectures using domain decomposition of the multi-block structured domain. The algorithm is applied to the nonlinear development of the tilt instability of a spheromak plasma configurations.

Nanosatellite maneuver planning for point cloud generation with a rangefinder

This paper discusses the application of a single beam laser rangefinder (LRF) to point cloud generation, shape detection, and shape reconstruction for a space-based space situational awareness (SSA) mission. The LRF is part of the payload of a chaser satellite tasked to image a resident space object (RSO). The one-dimensional (1D) nature of LRF returns significantly increases the complexity of the imaging task. To maximize coverage, a method to autonomously detect and fill gaps in sparse point cloud coverage using a narrow field of view (NFOV) camera in conjunction with the LRF is presented. First, relative orbital motion and scanning attitude motion are combined to generate a baseline 3D point cloud of the RSO. The effectiveness of pregenerated command profiles is analyzed by using a weighted edge reconstruction metric that scores how well a point cloud characterizes RSO shape. The design and characterization of multiple relative motion and attitude maneuver profiles, as well as the time and propellant cost of each profile, are presented with the assumption that the entire metrology chain is error free. Next, a three-part algorithm is used that creates a 3D panoramic map from stitched NFOV camera images, correlates the areas of sparse LRF coverage to the map, and generates attitude commands to close the coverage. This provides a consistent and reliable method to register positions of strike points relative to each other and to the NFOV image of the RSO with a priori knowledge of the RSO attitude. Gaps and sparse areas in LRF coverage are covered with strike points; the result is a point cloud of significantly higher resolution than that obtained with preprogrammed attitude profiles. The analysis bears particular relevance to power-constrained nanosatellite missions for space-based SSA for whom a multibeam LRF payload is not feasible. Maneuvers can now be designed on-line in real time; results presented validate the utility of a single-beam LRF as a tool for 3D imaging of RSO shapes.

Integrated Attitude and Orbit Control of an Interstellar Heliopause Probe

Solar sailing has been identified as a promising and enabling technology for future space missions; as such it is currently the object of a significant research effort within various space agencies, the academic world and industry. Active research and development activities have been performed by the European Space Agency (ESA) in the recent years to increase the technological readiness level of the key elements allowing the deployment and control of a spacecraft efficiently propelled by solar radiation pressure. A six-degree-of-freedom simulation environment has been developed to obtain the main results of the study and validate the attitude and orbit control concepts proposed for the Interstellar Heliopause Probe (IHP) mission is presented.

Accurate Modes Design of the Darwin Precursor Formation Flying Demonstration Mission

Abstract In preparation for planet detection ESA mIssIon Darwin, a range of demonstration activities allow to study the concepts and technology needed to achieve the required performance. The considered precursor mission is a fully operational demonstration mission that aims both at validating the overall metrology chain and manoeuvres feasibility and at performing imagery of science objects by aperture synthesis. This precursor mission consists in three spacecraft, two telescope flyers and one hub combiner, allowing to constitute an interferometric arm. To achieve science imaging requirements, the GNC nominal mode shall ensure Optical Path Difference accuracy of 23nm (1σ) and relative pointing accuracy of both telescopes line of sight of 62mas (1σ). This paper presents the GNC design to drive the constellation from coarse formation (based on RF and STR measurements) to the ultimate performance mode (scientific mode, based on fringe tracker OPD measurements and either ODL (Optical Delay Line) or µN FEEPs thrusters actuation).

Nonlinear Study of Spheromak Tilt Instability

The magnetohydrodynamic (MHD) tilt mode of a spheromak is thought to be the dominant unstable mode. In the present ^ study we attempt to apply our novel approximate Riemann solver for the threedimensional MHD equations to the analysis of the tilt instability of a spheromak contained in a prolate cylindrical vessel (flux conserver). We have applied an initial velocity perturbation to the spheromak and analyzed its evolution. The results are encouraging and the simulation shows that the spheromak tilts and evolves towards a minimum-energy state.

Pirarucu: The Mars moon prospector

The design of the Pirarucu mission has been undertaken by a team of undergraduates students at the Embry-Riddle Aeronautical University for the NASA/NIA 2015 Revolutionary Aerospace Systems Concepts - Academic Linkage (RASC-AL) competition. The team has completed a design iteration up to the equivalent of a mission concept review (MCR). Pirarucu is a Discovery class mission for prospecting Martian moons Phobos and Deimos. The concept is centered on a LADEE-like mothership that carries a set of twelve 12U Cube-Sats. The mothership is equipped with a suite of instruments similar to those on the Curiosity and the 2020 Mars rover, and is capable of performing laser spectroscopy, geological mapping, and chemical analysis of regolith samples. The mothership also serves as a telecommunication hub with the NASA Deep Space Network (DSN) and communicates directly with the CubeSats on a common data link (CDL) architecture implemented with L3 Communications Net-T technologies. The 12U CubeSats explore the surface of the moon and collect surface samples to be returned to the mothership for analysis. The CubeSats are split into two teams, one team can perform in-situ surface science and the other retrieves samples for analysis on the mothership. Both CubeSat teams can be refueled by the mothership and thus are able to perform multiple excursions to the moon for surface science and sample collection to ensure analysis of a variety of sites on the surfaces of the moons.

Micro-Propulsion System of a Thermospheric Explorer CubeSat

The major design driver for the Dipping Thermospheric Explorer (DipTE) CubeSat mission is that the satellite shall fly a significant amount of orbit arcs at altitudes of 300km and below. It is assumed that the DipTE satellite will be released in a circular orbit above the altitudes of scientific interest for the mission. A propulsion system will be employed to make the orbit elliptic and satisfy the design driver. The apogee of the elliptic orbit will be at the altitude of the initial circular orbit. The inclination of the orbit will stay the same. A few design iterations of the mission converged to a satellite configuration capable of storing 0.375 kg of propellant, and a propulsion system with a specific impulse Isp of 86 s. About 90 % of the total impulse of 313 Ns stored on board was budgeted to perform orbital maneuvers, with a 1 N thruster orbital maneuvering thruster (OMT). The remainder of 10% of the propellant mass has been budgeted for attitude control maneuvers, such as those performed during detumbling and initial attitude acquisition. The attitude control maneuvers are performed with the thrusters of a reaction control system (RCS). The 12 two-dimensional (2D) microthrusters of the RCS produce 40 mN each and are installed such that they provide three-axis control of the spacecraft. This paper describes the preliminary design of the propulsions system.© 2009 ASME