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Solutions to nonlinear optimal control problems (OCPs) exhibit dispersions under model uncertainties and it is desirable to generate optimal solutions that exhibit less sensitivity to model uncertainties. We propose a novel solution desensitization method dubbed “Reduced Desensitization Formulation (RDF)” by leveraging non-uniqueness of the solution of the costate differential equations when a hybrid indirect-direct optimization method is used. A key property of the RDF method is a significant reduction in the number of differential equations needed for generating desensitized solutions. This feature facilitates the generation of open-loop desensitized trajectories and makes the methodology applicable to OCPs with a larger number of uncertain parameters. To demonstrate the utility of the RDF method, three important classes of trajectory optimization problems are considered with uncertainty in the thrust magnitude of the propulsion system: (1) minimum-fuel low-thrust interplanetary rendezvous maneuvers, (2) low-thrust orbit-raising maneuvers, and (3) minimum-fuel high-thrust rocket-landing problems. For the considered problems with bang-bang control profiles, an analysis is presented on the change in the number of control switches between sensitive and desensitized optimal solutions. Numerical results demonstrate desensitization of the considered performance indices with respect to the thrust magnitude of the propulsion system. [ABSTRACT FROM AUTHOR]

A semi-analytical solution for circular-to-circular planar orbit transfers is presented. In particular, the problem is addressed with a judiciously chosen maneuver sequence consisting of radial thrust, velocity normal thrust, and coast arcs. The radial thrust segments admit fully-analytical solutions, while the velocity normal thrust portion only admits a semi-analytical solution. For a given constant-acceleration thrust value, the range of radii to which the orbit can be changed is presented with two different schemes. Orbit escape is also demonstrated with successive applications of the first scheme. However, all the presented solutions are suboptimal in terms of time and fuel use. [ABSTRACT FROM AUTHOR]

A new solar sail model that can be controlled passively using gravity stabilization and black-coating was designed. In this paper, a long boom with a tip-mass was used to realize gravity-gradient stabilization when a solar radiation pressure (SRP) was applied. The solar sail does not require extra controllers or fuel sources to maintain proper attitude for orbit raising schemes. In an orbit mission, attitude dynamics and control analysis of a solar sail are critical issues for successful performance. This paper investigated analytically and numerically the stability of 1-dof pitch motion of the solar sail model. Ignoring the SRP and conducting analytic interpretation, among the four equilibrium points of the solar sail, the solar sail was marginally stable in the standing attitude aligned along the direction of gravity or in the standing attitude upside down. The computation, including an SRP, showed that the state near two of the four equilibrium points (standing attitude and standing attitude upside down) were oscillating within a certain boundary if the perturbation was small and stable. The validation was verified by checking the phase portrait and the power spectral density. The simulation of 3-dof orbit demonstration mission of the presented model with translation motions shows the result of increased orbital altitude change, implying the feasibility of mission performance. [ABSTRACT FROM AUTHOR]

This paper presents novel design concepts and a robust sliding mode control (RSMC) algorithm for a solar sail optimized for an Earth-orbit-raising demonstration mission. The design of the solar sail is focused on minimizing mission difficulty and control effort during the Earth-orbit-rising mission. To this end, a gravity-stabilized long boom and one-sided black-coating film are introduced in the solar sail system, and a reaction wheel is used as a control torque actuator. The strategic combination of these components modifies the attitude control task from 90° back-and-forth pitch motion control for maintaining the proper attitude for receiving solar radiation pressure (SRP) to that of stabilizing the sail attitude perturbed from gravity-stabilized equilibrium. This simplification of the control task can minimize the mission difficulty and control input energy because attitude control is required only for small deviations from the attitude equilibrium, while the one-sided black-coated film with a reflective surface on the opposite side ensures that SRP propulsion effectively pushes the sail toward the orbit-rising direction. Thus, the control torque of the reaction wheel size can be reduced. To validate the Earth-orbit-raising mission performance, a quaternion-based RSMC approach is established for the six degree-of-freedom dynamic model. The simulation results confirm that the proposed solar sail system can stabilize its attitude with minimal control torque energy of the reaction wheels during Earth-orbit-raising missions. [ABSTRACT FROM AUTHOR]

In this study, the semi-optimal attitude steering logic for low-thrust orbit raising considering both the power generation and the angular momentum constraints is analytically derived to minimize the transfer time. In the proposed method, the attitude steering is realized by the rotation around the sun direction and the successive rotation around the rotation axis of solar array driving mechanism (SADM). These two rotation angles are determined so that the relative attitude error with respect to the ideal attitude is minimized in the presence of the angular momentum and torque constraints. The numerical simulations show that the degradation of the transfer time by the proposed logic is only less than several percent compared to the case without the angular momentum and torque constraints. [ABSTRACT FROM AUTHOR]

This study considers the minimum time orbit-raising problem of geostationary spacecraft with low-propulsion thrusters. This problem is equivalent to determining an appropriate thrust direction during orbit raising. This study proposes a closedloop thruster steering law that determines the thrust direction based on the optimal feedback gains and control errors of each orbital element. The feedback gains of the steering law are assumed to be the functions of orbital elements and are optimized by a meta-heuristic method. The orbital semi-major axis, eccentricity, and inclination are considered as independent variables for expressing the gains. Numerical simulations show that whichever orbital elements is selected as an independent variable, the same performances are obtained. Therefore, regardless of the initial orbital elements, by selecting the independent variable appropriately, the proposed method can always solve the minimum time orbit-transfer problem. [ABSTRACT FROM AUTHOR]

In its typical arrangement, an Electric Solar Wind Sail (or E-sail) consists of a grid of long tethers charged by means of a high-voltage source that creates an electrostatic sheath, while an electron emitter is used to remove trapped electrons. Such a grid of conducting tethers enables the generation of a thrust by deviating the charged particles coming from the Sun with which it interacts electrostatically. However, the difficulty of deploying and maintaining stretched such large structures has been leading the scientific community to consider simpler configurations, such as the one in which the E-sail consists of a single spinning tether. This Short Communication fits into this context, by first describing the thrust and torque vectors of a single-tether E-sail. A drawback of the single-tether arrangement is that it is more difficult to maneuver than a multi-tether configuration, because a modulation of the electric voltage, which is one possible means of controlling the attitude of a multi-tether E-sail, would not enable the generation of any control torque. For this reason, this work also proposes a guidance scheme for orbit raising or lowering scenarios, in which the tether is alternately charged and discharged while the inertial direction of the spacecraft spin axis always remains unchanged. [ABSTRACT FROM AUTHOR]

The FORMOSAT-3/COSMIC spacecraft constellation consisting of six LEO satellites is the world's first operational GPS Radio Occultation (RO) mission. The mission is jointly developed by Taiwan's National Space Organization (NSPO) and the United States' UCAR in collaboration with NSF, USAF, NOAA, NASA, NASA's Jet Propulsion Laboratory, and the US Naval Research Laboratory. The FORMOSAT-3/COSMIC satellites were successfully launched from Vandenberg US AFB in California at 0140 UTC 15 April 2006 into the same orbit plane of the designated 516 km altitude. The mission goal is to deploy the six satellites into six orbit planes at 800 km altitude with a 30-degree separation for evenly distributed global coverage. All six FORMOSAT-3/COSMIC satellites are currently maintaining a satisfactory good state-of-health. Five out of six satellites have reached their final mission orbit of 800 km as of November 2007. The data as received by FORMOSAT-3/COSMIC satellites constellation have been processed in near real time into 2500 good ionospheric profiles and 1800 good atmospheric profiles per day. These have outnumbered the worldwide radiosondes (∼900 mostly over land) launched from the ground per day. The processed atmospheric RO data have been assimilated into the Numerical Weather Prediction (NWP) models for real-time weather prediction and typhoon/hurricane forecasting by many major weather centers in the world. This paper describes the FORMOSAT-3/COSMIC satellite constellation system performance and the mission results that span the period from April 2006 to October 2007; and reviews the prospect of a future follow-on mission. [ABSTRACT FROM AUTHOR]

The current paper describes the detailed operational planning process required for the decommissioning of the Inmarsat 2F3 satellite. The background to events which lead to the decision to decommission this satellite is presented and descriptions of the methods used to monitor the onboard propellant reserves are included in order to explain their role in determining the date of service termination. Orbital and operational factors are elaborated and then folded into a baseline plan for executing a sequence of orbit raising manoeuvres that would be necessary to increase the satellite altitude above the busy geostationary arc. Subsystem passivation activities to be performed on arrival at the final graveyard orbit are discussed and the paper concludes with the story of what actually happened with the first ever Inmarsat satellite decommissioning. [ABSTRACT FROM AUTHOR]