The Experts below are selected from a list of 120 Experts worldwide ranked by ideXlab platform
Lorenzo Casalino - One of the best experts on this subject based on the ideXlab platform.
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High-accuracy optimal finite-thrust trajectories for Moon escape
Acta Astronautica, 2017Co-Authors: Hong Xin Shen, Lorenzo CasalinoAbstract:Abstract The optimization problem of fuel-optimal trajectories from a low circular Moon orbit to a target Hyperbolic Excess Velocity vector using finite-thrust propulsion is solved. The ability to obtain the most accurate satisfaction of necessary optimality conditions in a high-accuracy dynamic model is the main motivation of the current study. The solutions allow attaining anytime-return Earth-interface conditions from a low lunar orbit. Gravitational effects of the Sun, Earth, and Moon are included throughout the entire trajectory. Severe constraints on the fuel budget combined with high-accuracy demands on the endpoint conditions necessitate a high-fidelity solution to the trajectory optimization problem and JPL DE405 ephemeris model is used to determine the perturbing bodies' positions. The optimization problem is solved using an indirect method. The optimality of the solution is verified by an application of Pontryagin's maximum principle. More accurate and fuel-efficient trajectories are found for the same mission objectives and constraints published in other research, emphasizing the advantages of this technique. It is also shown that the thrust structure consists of three finite burns. In contrast to previous research, no singular arc is required in the optimal solutions, and all the controls appear bang-bang.
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Optimal Low-Thrust Escape Trajectories Using Gravity Assist
Journal of Guidance Control and Dynamics, 1999Co-Authors: Lorenzo Casalino, Guido Colasurdo, D. PastroneAbstract:Electric propulsion provides a spacecraft with continuous steering capabilities, which can be used to approach a planet with Hyperbolic Excess Velocity that enhances the gravity assist. Low-thrust trajectories to escape from the solar system are considered in the present paper, which searches for the strategy that maximizes the spacecraft energy for assigned payload and engine operating time. The optimal conditions to escape using electric propulsion and gravity assist are presented for the cases of free-height and minimum-height ybys. Optimal trajectories that exploit Jupiter or Venus ybys have been computed for constant exhaust power with either constant or variable speci c impulse; the procedure is also able to determine the optimal power level and to suggest when it is convenient to switch the engine on and off. The bene t that system performance can receive by increasing the number of controls, i.e., by adding the possibility of coast arcs and engine throttling to the thrust direction control, is also noted.
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Simple Strategy for Powered Swingby
Journal of Guidance Control and Dynamics, 1999Co-Authors: Lorenzo Casalino, Guido Colasurdo, D. PastroneAbstract:The use of thrust during a close encounter with a celestial body can signi cantly improve the ef ciency of the swingby maneuver. A parametric analysis of a powered swingby with the moon is carried out to nd the optimal combined use of a gravity assist and propulsion. A single impulse, usually applied inside the satellite’s sphere of in uence, is minimized to achieve an assigned energy with respect to the Earth, for an assigned Hyperbolic Excess Velocity on entering the sphere. Numerical results are presented, but emphasis is placed on the physical reasoning of the suggested strategies that are sometimes different from those found in recent literature. A longer stay on the low-energy, large-de ection hyperbola is always convenient. When geocentric energy is increased, thrust is applied after the passage at the periapsis, with a small misalignment with respect to the spacecraft Velocity, to obtain a larger turn angle from the swingby. If reduction of the geocentric energy is required, the selenocentric energy is increased or reduced, depending on the magnitude of the inbound Hyperbolic Excess Velocity, and the corresponding strategies are rather different.
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Optimal low-thrust trajectories using flyby
Advances in the Astronautical Sciences, 1998Co-Authors: Lorenzo Casalino, Guido Colasurdo, D. PastroneAbstract:Electric propulsion provides a spacecraft with continuous steering capabilities which can be used to approach a planet with the Hyperbolic Excess Velocity that enhances the gravity assist. Low-thrust trajectories to escape from the solar system are considered in the present paper; the strategy that maximizes the spacecraft energy for assigned payload and engine operating time is searched for. The optimal conditions to escape using electric propulsion and gravity assist are presented, in the cases of free-height and minimum-height flybys. Optimal trajectories that exploit Jupiter or Venus flybys have been computed for constant exhaust power with either constant or variable specific impulse; the procedure is also able to determine the optimal power level and to suggest when it is convenient to switch the engine on and off. The benefit that system performance can receive by increasing the number of controls, i.e., by adding the possibility of coast arcs and engine throttling to the thrust direction control, is also pointed out.
D. Pastrone - One of the best experts on this subject based on the ideXlab platform.
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Optimal Low-Thrust Escape Trajectories Using Gravity Assist
Journal of Guidance Control and Dynamics, 1999Co-Authors: Lorenzo Casalino, Guido Colasurdo, D. PastroneAbstract:Electric propulsion provides a spacecraft with continuous steering capabilities, which can be used to approach a planet with Hyperbolic Excess Velocity that enhances the gravity assist. Low-thrust trajectories to escape from the solar system are considered in the present paper, which searches for the strategy that maximizes the spacecraft energy for assigned payload and engine operating time. The optimal conditions to escape using electric propulsion and gravity assist are presented for the cases of free-height and minimum-height ybys. Optimal trajectories that exploit Jupiter or Venus ybys have been computed for constant exhaust power with either constant or variable speci c impulse; the procedure is also able to determine the optimal power level and to suggest when it is convenient to switch the engine on and off. The bene t that system performance can receive by increasing the number of controls, i.e., by adding the possibility of coast arcs and engine throttling to the thrust direction control, is also noted.
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Simple Strategy for Powered Swingby
Journal of Guidance Control and Dynamics, 1999Co-Authors: Lorenzo Casalino, Guido Colasurdo, D. PastroneAbstract:The use of thrust during a close encounter with a celestial body can signi cantly improve the ef ciency of the swingby maneuver. A parametric analysis of a powered swingby with the moon is carried out to nd the optimal combined use of a gravity assist and propulsion. A single impulse, usually applied inside the satellite’s sphere of in uence, is minimized to achieve an assigned energy with respect to the Earth, for an assigned Hyperbolic Excess Velocity on entering the sphere. Numerical results are presented, but emphasis is placed on the physical reasoning of the suggested strategies that are sometimes different from those found in recent literature. A longer stay on the low-energy, large-de ection hyperbola is always convenient. When geocentric energy is increased, thrust is applied after the passage at the periapsis, with a small misalignment with respect to the spacecraft Velocity, to obtain a larger turn angle from the swingby. If reduction of the geocentric energy is required, the selenocentric energy is increased or reduced, depending on the magnitude of the inbound Hyperbolic Excess Velocity, and the corresponding strategies are rather different.
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Optimal low-thrust trajectories using flyby
Advances in the Astronautical Sciences, 1998Co-Authors: Lorenzo Casalino, Guido Colasurdo, D. PastroneAbstract:Electric propulsion provides a spacecraft with continuous steering capabilities which can be used to approach a planet with the Hyperbolic Excess Velocity that enhances the gravity assist. Low-thrust trajectories to escape from the solar system are considered in the present paper; the strategy that maximizes the spacecraft energy for assigned payload and engine operating time is searched for. The optimal conditions to escape using electric propulsion and gravity assist are presented, in the cases of free-height and minimum-height flybys. Optimal trajectories that exploit Jupiter or Venus flybys have been computed for constant exhaust power with either constant or variable specific impulse; the procedure is also able to determine the optimal power level and to suggest when it is convenient to switch the engine on and off. The benefit that system performance can receive by increasing the number of controls, i.e., by adding the possibility of coast arcs and engine throttling to the thrust direction control, is also pointed out.
Pekka Janhunen - One of the best experts on this subject based on the ideXlab platform.
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Optimal Interplanetary Rendezvous Combining Electric Sail and High Thrust Propulsion System
Acta Astronautica, 2011Co-Authors: Alessandro Antonio Quarta, Giovanni Mengali, Pekka JanhunenAbstract:Abstract The aim of this paper is to study, from a mission analysis point of view, the performance of a hybrid propulsion concept for a two-dimensional transfer towards a planet of the Solar System. The propulsion system is obtained by combining a chemical thruster, used for the phases of Earth escape and/or target planet capture, with an electric sail, which provides a continuous thrust during the heliocentric transfer. Two possible mission scenarios are investigated: in the first case the sailcraft reaches the target planet with zero Hyperbolic Excess Velocity, thus performing a classical rendezvous mission in a heliocentric framework. In the second mission scenario, a given final Hyperbolic Excess Velocity relative to the planet is tolerated in order to decrease the mission flight time. The amount of final Hyperbolic Excess Velocity is used as a simulation parameter for a tradeoff study in which the minimum flight time is related to the total Velocity variation required by the chemical thruster to accomplish the mission, that is, for Earth escape and planetary capture.
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Optimal interplanetary rendezvous combining electric sail and high thrust propulsion system
Acta Astronautica, 2011Co-Authors: Alessandro Antonio Quarta, Giovanni Mengali, Pekka JanhunenAbstract:The aim of this paper is to study, from a mission analysis point of view, the performance of a hybrid propulsion concept for a two-dimensional transfer towards a planet of the Solar System. The propulsion system is obtained by combining a chemical thruster, used for the phases of Earth escape and/or target planet capture, with an electric sail, which provides a continuous thrust during the heliocentric transfer. Two possible mission scenarios are investigated: in the first case the sailcraft reaches the target planet with zero Hyperbolic Excess Velocity, thus performing a classical rendezvous mission in a heliocentric framework. In the second mission scenario, a given final Hyperbolic Excess Velocity relative to the planet is tolerated in order to decrease the mission flight time. The amount of final Hyperbolic Excess Velocity is used as a simulation parameter for a tradeoff study in which the minimum flight time is related to the total Velocity variation required by the chemical thruster to accomplish the mission, that is, for Earth escape and planetary capture. © 2010 Elsevier Ltd. All rights reserved.
M. S. Konstantinov - One of the best experts on this subject based on the ideXlab platform.
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Comparative Analysis of the Effectiveness of Methods for Solving Boundary Value Problems of the Maximum Principle for Low Thrust Trajectory Optimization
2018Co-Authors: M. S. Konstantinov, Min TheinAbstract:The effectiveness of existing methods for minimizing the function in the finite-dimensional vector space for solving boundary value problems (BVP) of the maximum principle is analyzed. The two formulations of the low thrust trajectory optimization problem to Mars are considered. The first formulation assumes the possibility of an ideal regulation of the thrust and specific impulse of the electric propulsion (ideally regulated engine). In the second, it is assumed that the magnitudes of the thrust and the specific impulse are constant in the active sections and the electric propulsion (EP) can be switched on and off repeatedly (unregulated engine). The Pontryagin maximum principle which allowing reducing the optimal control problem to the BVP for the system of ordinary differential equations is used for both formulations. The BVP is two-point and its order is seven. The solution of the BVP is formulated as the minimization problem for the sum of squares of the BVP residuals. It is required to find the solution in which the minimum found is zero. In the event that there are several such minima, then it is proposed to choose the solution of the BVP in which the final mass of the spacecraft turns out to be the maximum. In both problem formulations, the launch date (April 20, 2035) and the flight time to the vicinity of Mars (350 days) are considered as given. The Hyperbolic Excess Velocity at the start from the Earth and the Hyperbolic Excess Velocity when approaching to Mars are equal zero. The initial mass of the spacecraft is equal to 1000 kg. The maximum reactive power of the ideally regulated EP is assumed equal to 5.88 kW. The thrust and specific impulse of the unregulated EP are assumed to be 0.4 N and 3000 s respectively. The optimization criterion is the final mass of spacecraft (mass of spacecraft at the moment of SC approach to the Mars vicinity). This mass is maximized. Effectiveness of different types of local, global and hybrid methods (trust-region-dogleg, trust-region-reflective, levenberg-marquardt, particle swarm optimization, genetic algorithm, simulated annealing, hybrid of particle swarm optimization and active – set algorithm, hybrid of genetic algorithm and sequential quadratic programming, hybrid simulated annealing and active – set algorithm, CMAES (Covariance Matrix Adaptation Evolution Strategy)) for solving BVP is analyzed. The indicators of the effectiveness of the method were considered as: 1) whether the solution is found that ensures the zero value of the sum of the squares of the BVP residuals; 2) which of the local maxima of the final mass of spacecraft is obtained (local or global); 3) the number of calculations of the sum of the squares of the BVP residuals in the iteration process. Seven sets of zeros, plus one and minus one were considered as the initial approximation for adjoint variables at the initial point of the heliocentric trajectory. For each set of initial approximation, the BVP was solved by using the tested methods. The use of CMAES for all considered sets of initial approximation made it possible to obtain the solution of the BVP and ultimately obtain an absolute maximum of the final mass of the spacecraft. The first three listed methods did not provide such an opportunity. It should be noted that using CMAES requires a relatively large number of iterations. In the case, the levenberg-marquardt method turned out to be more efficient than CMAES because the number of calculations of the minimized sum of squares of residuals in 100 times less than when using CMAES. But let repeat, the method of levenberg-marquardt did not always give the solution of the problem. The hybrid algorithms based on particles swarm optimization, the genetic algorithm proved to be quite good. As a rule, they allowed obtaining the optimal solution. But the number of iterations was more (significantly more) than required by CMAES. The general conclusion from the analysis is as follow: The use of CMAES at least at the first stage of the solution of the BVP of the maximum principle is expedient. To accelerate the iterative procedure, it is possible to use the method based on local search (Newton methods, the conjugate gradient method, the levenberg-marquardt method) at the second stage of the investigation (after finding the global extreme region).
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Method of interplanetary trajectory optimization for the spacecraft with low thrust and swing-bys
Acta Astronautica, 2017Co-Authors: M. S. Konstantinov, M. TheinAbstract:Abstract The method developed to avoid the complexity of solving the multipoint boundary value problem while optimizing interplanetary trajectories of the spacecraft with electric propulsion and a sequence of swing-bys is presented in the paper. This method is based on the use of the preliminary problem solutions for the impulsive trajectories. The preliminary problem analyzed at the first stage of the study is formulated so that the analysis and optimization of a particular flight path is considered as the unconstrained minimum in the space of the selectable parameters. The existing methods can effectively solve this problem and make it possible to identify rational flight paths (the sequence of swing-bys) to receive the initial approximation for the main characteristics of the flight path (dates, values of the Hyperbolic Excess Velocity, etc.). These characteristics can be used to optimize the trajectory of the spacecraft with electric propulsion. The special feature of the work is the introduction of the second (intermediate) stage of the research. At this stage some characteristics of the analyzed flight path (e.g. dates of swing-bys) are fixed and the problem is formulated so that the trajectory of the spacecraft with electric propulsion is optimized on selected sites of the flight path. The end-to-end optimization is carried out at the third (final) stage of the research. The distinctive feature of this stage is the analysis of the full set of optimal conditions for the considered flight path. The analysis of the characteristics of the optimal flight trajectories to Jupiter with Earth, Venus and Mars swing-bys for the spacecraft with electric propulsion are presented. The paper shows that the spacecraft weighing more than 7150 kg can be delivered into the vicinity of Jupiter along the trajectory with two Earth swing-bys by use of the space transportation system based on the "Angara A5" rocket launcher, the chemical upper stage "KVTK" and the electric propulsion system with input electrical power of 100 kW.
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Estimation of an opportunity of Mercury mission with use of solar electric propulsion
Acta Astronautica, 2002Co-Authors: M. S. Konstantinov, Gennadiy FedotovAbstract:Abstract Estimation of transport opportunities of delivery of a research spacecraft (SC) to Mercury with the use of launcher “Proton”, chemical upper stage (ChUS) as “D” and solar electric propulsion upper stage (SEPS) on the basis of thrusters ESA-XX and SPT is analyzed. It is assumed that Russian launch vehicle “Proton” and upper state “Block D” deliver a spacecraft into the geocentric Hyperbolic orbit. Then “Block D” is separated. The heliocentric arc of the trajectory and operations in the vicinity of Mercury are supported by SEPS. It is supposed, that the change of SEPS thrust value according to available electrical power occurs due to change in propellant (Xenon) flow rate. The problem was formulated as follows. For the known characteristics of the ChUS, the SC and its systems are required to determine the unknown design SC parameters, circuit of transfer and control program of a thrust vector, ensuring delivery in the vicinity of Mercury of maximum SC mass at a given transfer duration. Chosen parameters and control functions are: mass of ChUS propellant (or value of Hyperbolic Excess Velocity connected with it), direction of Hyperbolic Excess Velocity, launch date, the control program of the thrust vector of the ion-engine (orientation of a thrust vector, moments of the engine firings and shut-down).
C Ocampo - One of the best experts on this subject based on the ideXlab platform.
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optimization of impulsive trajectories from circular orbit to an Excess Velocity vector
Journal of Guidance Control and Dynamics, 2012Co-Authors: Drew R Jones, C OcampoAbstract:Feasible transfer trajectories are constructed to serve as initial guesses for determining constrained optimal impulsive escape trajectories, from a circular orbit to a target Hyperbolic Excess Velocity vector. The proximity of these feasible solutions to their corresponding optima are quantified. The objective of this work is to improve the feasible solutions, thereby enabling general anytime escape trajectories to be targeted that result in substantially reduced fuel expenditure. The procedure is currently restricted to one- and three-impulse escape trajectories from a circular parking orbit. Two separate three-impulse methods are presented, each specific to whether the time of flight is free or fixed. Numerical results, obtained using nonlinear programming algorithms, are presented to validate the robustness of the method. Analysis and results are nondimensional, and therefore are applicable to departures from a circular orbit about any celestial body.
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optimal impulsive escape trajectories from a circular orbit to a Hyperbolic Excess Velocity vector
AIAA AAS Astrodynamics Specialist Conference, 2010Co-Authors: C OcampoAbstract:Feasible impulsive transfers from a circular parking orbit to some Hyperbolic Excess Velocity vector are constructed, and used as initial guesses in solving corresponding, minimum V , constrained parameter optimization problems. Much attention is devoted to quantifying and enhancing the proximity of the feasible escape to the optimal, thereby improving convergence. The impulsive solutions may be further utilized as initial guesses for a procedure to determine optimal, nite-burn, continuous transfers between celestial bodies. Signicant fuel costs associated with certain escape geometries motivates in-depth understanding of this fundamental optimization problem. Reducing these costs provides increased: landing site coverage, abort capability, and parking orbit geometries which may have otherwise not met mission constraints. 1-impulse and 3-impulse (time-of-ight free and