Formation flying mission for the UW Dawgstar satellite

Abstract-An overview of a small satellite (< 15 kg) being designed and built by the University of Washington for multiple small satellite formation flying is presented. The Dawgstar nanosatellite is one of three satellites being built for a three satellite formation experiment termed ION-F, and is unique in its propulsive capability. The satellites will also utilize an integrated GPS/cross-link system to allow fast and accurate update of relative satellite positions. The three satellites are each a part of the AFRL/DARPA/NASA University Nanosatellite program, which addresses building unique small satellites and developing coordination experiments for space. The satellites are being designed for a Space Shuttle launch in January 2002

Propulsion Trade Studies for Spacecraft Swarm Mission Design

Spacecraft swarms constitute a challenge from an orbital mechanics standpoint. Traditional mission design involves the application of methodical processes where predefined maneuvers for an individual spacecraft are planned in advance. This approach does not scale to spacecraft swarms consisting of many satellites orbiting in close proximity; non-deterministic maneuvers cannot be preplanned due to the large number of units and the uncertainties associated with their differential deployment and orbital motion. For autonomous small sat swarms in LEO, we investigate two approaches for controlling the relative motion of a swarm. The first method involves modified miniature phasing maneuvers, where maneuvers are prescribed that cancel the differential delta V of each CubeSat's deployment vector. The second method relies on artificial potential functions (APFs) to contain the spacecraft within a volumetric boundary and avoid collisions. Performance results and required delta V budgets are summarized, indicating that each method has advantages and drawbacks for particular applications. The mini phasing maneuvers are more predictable and sustainable. The APF approach provides a more responsive and distributed performance, but at considerable propellant cost. After considering current state of the art CubeSat propulsion systems, we conclude that the first approach is feasible, but the modified APF method of requires too much control authority to be enabled by current propulsion systems

Distributed estimation and control technologies for formation flying spacecraft by Philip Andrew Ferguson.

Thesis (S.M.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2003.Includes bibliographical references (p. 115-120).S.M

Development and testing of model predictive control strategies for spacecraft formation flying

Satellite Formation Flying (SFF) is a key technology for several future missions, since, with respect to a single spacecraft, it allows better performances, new capabilities, more flexibility and robustness to failure and cost reduction. Despite these benefits, however, this new concept poses several signicant design challenges and requires new technologies. The Guidance, Navigation and Control (GNC) system is a key element in the SFF concept since it must be reliable in coordinating all the satellites fying in formation during each mission phase, guaranteeing formation integrity and preventing from formation evaporation, and, at the same time, efficient in using the limited on board resources. Model Predictive Control (MPC), also referred to as Receding Horizon Control, is a modern optimal control technique that seems to be suitable for these purposes because of its three main features: model-based control scheme, constraints handling ability and replanning nature. The final aim of my Ph.D. activities was to develop and test MPC strategies for SFF applications. This task was accomplished by means of both computer simulations and experimental tests conducted on both the MIT Synchronized Position Hold Engage & Reorient Experimental Satellites (SPHERES) testbed and the SFF Hardware Simulator under development at the Center of Studies and Activities for Space "Giuseppe Colombo" (CISAS), University of Padova. MPC capabilities were first tested in computer simulations in carrying out a formation acquisition maneuver for two space vehicles, taking into account two scenarios: a Leader-Follower (LF) formation and Projected Circular Orbit (PCO) formation. The performances of the MPC-based controller were compared with those of a Linear Quadratic Regulator (LQR) based controller in the presence of active constraints on the maximum control acceleration, evaluating also the effects of the gravitational harmonics J2 and J3 and atmospheric drag perturbations on the proposed maneuvers. Simulation results of both scenarios showed that, with similar performances in tracking the same reference state trajectory in terms of settling time, the MPC controller is more efficient (less delta-v requirement) than the LQR controller also in the perturbed cases, allowing a delta-v requirement reduction by 40% in the LF formation scenario and by 30% in the PCO formation scenario. The next activity concerned the development of some guidance and control strategies for a Collision-Avoidance scenario in which a free-flying chief spacecraft follows temporary off-nominal conditions and a controlled deputy spacecraft performs a collision avoidance maneuver. The proposed strategy consists on a first Separation Guidance that, using a computationally simple, deterministic and closed-form algorithm, takes charge of avoiding a predicted collision. When some safe conditions on the relative state vector (position and velocity) are met, a subsequent Nominal Guidance takes over. Genetic Algorithms are used to compute a pair of reference state trajectories in order to place the deputy spacecraft in a bounded safe or "parking" trajectory, while minimizing the propellant consumption and avoiding the formation evaporation. The performances of a LQR and a MPC in tracking these reference trajectories were compared, showing how a MPC controller can reduces the total delta-v requirement by 5 - 10% with respect to a LQR controller. MPC capabilities were then evaluated on the MIT SPHERES testbed in simulating the close-proximity phase of the rendez-vous and capture maneuver for the Mars Orbital Sample Return (MOSR) scenario. Better performances of MPC with respect to PD in executing this maneuver were conrmed both in a Matlab simulator and in the MIT SPHERES software simulator, with a total delta-v requirement reduction by 10-15 %. The proposed MPC control strategy was then tested using the SPHERES Flat Floor facility at the MIT Space System Laboratory. The last part of my research activities was devoted to the SFF Hardware Simulator of the University of Padova. My contributions to this project dealt with: (a) conclusion of the designing, building and testing of the five main subsystems of the hardware simulator; (b) software development for the hardware simulator and its Matlab software simulator; (c) preparatory experimental activities aimed at characterizing the thrust force performed by the on board thrusters and estimating the hardware simulator inertia properties; and (d) test of attitude control maneuvers with the use of predictive controllers. In particular, three main tests were carried out with the hardware simulator moving at one degree of freedom about the yaw axis. The first one aimed at tuning a Kalman Filter to properly estimate the yaw axis angular velocity using a double-integrator as dynamic model and angular position measurements provided by the yaw quadrature encoder. With the use of a simple Kalman Filter, the yaw angular position and velocity could be estimated with an error less than 0.1 ° and 0.1°/s, respectively. In the second test, an explicit MPC was used to perform a 170° slew maneuver of the hardware simulator attitude module about the yaw axis. The final target angular position was reached with an error less than 0.5° in 20 s. In the third test, a 3 degrees of freedom attitude reference trajectory was first computed using pseudospectral optimization methods for a repointing maneuver with active constraints on the attitude trajectory. The state trajectory was then projected along the satellite z-Body axis and tracked in the hardware simulator using an explicit MPC. Experimental results showed that with an explicit MPC the reference trajectories can be tracked with an error less that 1.5° for the angular position and less than 1°/s for the angular velocity, both in dynamic conditions. The final target state was reached with an error less than the estimation accuracy. The SFF Hardware Simulator is a ground-based testbed for the development and verification of GNC algorithms that in the present configuration allows the development and testing of advanced controls for attitude motion and in its final form will enable the derivation of control strategies for Formation Flight and Automated Rendezvous and Docking

Design of a thruster-assisted control for 1U CubeSat orbit maintenance and drag mitigation with NASA 42

In the last ten years, CubeSats have gained popularity in the context of spatial missions in low Earth orbits (LEO). They have a simple construction and complexity and a great variety of possibilities at an affordable cost. This makes them an esteemed option for universities and research groups that want to test new technologies in space at a reduced cost. Therefore, this thesis aims to design a control system capable of controlling the behavior of a 1U CubeSat to solve the main problem of these types of satellites, which is the short duration of their lifetime. Due to the atmospheric density in LEO, drag is the principal cause of impeding prolonged missions. So the best way to solve it is to implement a thruster capable of mitigating the drag to maintain the altitude. Due to the difficulty of testing the thruster in real life, it is selected a software to simulate the behavior of the CubeSat in space. The NASA 42 Simulator is an open-source software integrated into the OpenSatKit environment that emulates accurately the spacecraft attitude, orbit dynamics, and control. Moreover, it is necessary to do a feasibility study to analyze if the project can be carried out in real life. This study analyses which is the best thruster to implement in the 1U CubeSat. Also, it focuses on the duration of the fuel of the satellite and compares it to the lifespan of the spacecraft when orbits without the thruster. After executing the simulations, the CubeSat’s orbit decays in approximately two years without implementing the thruster. On the other hand, when the thruster is activated, the mission reaches seven and a half years. With this outcome, it can be concluded that the designed control system can be conducted in real life

Parallel orbit propagation and the analysis of satellite constellations

Thesis (M.S.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 1995.Includes bibliographical references (p. 259-267).by Scott Thoams Wallace.M.S

1999 Flight Mechanics Symposium

This conference publication includes papers and abstracts presented at the Flight Mechanics Symposium held on May 18-20, 1999. Sponsored by the Guidance, Navigation and Control Center of Goddard Space Flight Center, this symposium featured technical papers on a wide range of issues related to orbit-attitude prediction, determination, and control; attitude sensor calibration; attitude determination error analysis; attitude dynamics; and orbit decay and maneuver strategy. Government, industry, and the academic community participated in the preparation and presentation of these papers

Modeling Satellite Formations In The Presence Of Perturbations

The potential benefits of autonomous satellite formation flying in such areas as high- resolution remote sensing, and sparse aperture radar, has stimulated interest in modeling the satellite environment for feasibility and simulation studies to help explore and define the technical challenges that must be solved in order to achieve successful autonomous satellite formations. The purpose of this paper is to develop and describe a numerical simulation of the orbital environment including central force field perturbations and atmospheric drag effects which will be a useful analytical tool for investigating issues relating to maintaining satellite formations in low-earth-orbit. Many of the studies done in this area confine their research to circular orbits, with and without perturbation effects. This study will investigate apply orbital dynamic equations to the problem of maintaining satellite formations in both circular and elliptical orbits, with and without the presence of J2 gravity perturbation effects and atmospheric drag. This effort is primarily focused on modeling the orbital mechanics of one and two satellites in the presence of J2 and drag perturbations This effort is being performed as part of a multi-disciplined University of Central Florida KnightSat project, sponsored by the Air Force, to develop a two-satellite formation in the nanosatellite class, for investigating issues related to using formation satellites for remote earth sensing, to develop three-dimensional mapping

Attitude and formation control design and system simulation for a three-satellite CubeSat mission

Thesis (S.M.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2013.This electronic version was submitted and approved by the author's academic department as part of an electronic thesis pilot project. The certified thesis is available in the Institute Archives and Special Collections.Cataloged from department-submitted PDF version of thesisIncludes bibliographical references (p. 113-115).Spacecraft formation flight has been identified as a critical enabling technology for achieving many scientific, commercial, and military objectives. One of the primary challenges of a formation flight mission is the control of the relative motion between spacecraft. Before any flagship missions will launch, technology development missions will be required to demonstrate the utility and functionality of formation flying systems. This thesis describes the complete attitude and formation control design for the MotherCube formation flight technology demonstration mission in LEO. A model of the spacecraft's sensors and actuators is developed and analyzed. Using curvilinear orbit theory, a simple LQR control law is used to generate a set of desired relative accelerations for formation control. A newly developed two-tier numerical allocation scheme is used alongside an independent PD attitude control law to generate a set of actuator commands which provides 3-axis attitude stabilization as well as formation control with guaranteed feasibility of actuator commands. An Extended Kalman Filter was developed to estimate the system attitude and angular rate from sensor measurements. To test these algorithms, a simulation environment was developed. This environment includes realistic models of space environment and the major perturbation effects which a LEO spacecraft formation would encounter. In order to improve the fidelity, a new intermediate-accuracy method for computing attitude-dependent aerodynamic and solar effects was also developed. Finally, results from the simulation are used numerically validate the dual-allocator approach, assess the performance of the control laws and provide system level metrics such as fuel use and required maneuver time.by Austin Kyle Nicholas.S.M