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Trajectory optimization of low-thrus...

Graham, Kathryn Frances.

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Trajectory optimization of low-thrust orbit transfers.

紀錄類型:	書目-電子資源 : Monograph/item
正題名/作者:	Trajectory optimization of low-thrust orbit transfers./
作者:	Graham, Kathryn Frances.
出版者:	Ann Arbor : ProQuest Dissertations & Theses, : 2015,
面頁冊數:	153 p.
附註:	Source: Dissertation Abstracts International, Volume: 77-12(E), Section: B.
Contained By:	Dissertation Abstracts International77-12B(E).
標題:	Aerospace engineering. -
電子資源:	http://pqdd.sinica.edu.tw/twdaoapp/servlet/advanced?query=10154476
ISBN:	9781369095692

Trajectory optimization of low-thrust orbit transfers.
Graham, Kathryn Frances.

Trajectory optimization of low-thrust orbit transfers. - Ann Arbor : ProQuest Dissertations & Theses, 2015 - 153 p.

Source: Dissertation Abstracts International, Volume: 77-12(E), Section: B.

Thesis (Ph.D.)--University of Florida, 2015.

Since the world's first artificial satellite, the Sputnik 1, completed its first orbit around the Earth, we have persistently developed new technology with an emphasis on cost-effective ways to place objects in space. NASA has identified low-thrust propulsion as a critical technology that would greatly reduce the cost of orbital insertion of spacecraft. To date, low-thrust propulsion has been studied in extensive detail and used mostly for interplanetary missions. Little attention, however, has been applied to the usage of low-thrust technology for orbital insertion. This research seeks to devise approaches for obtaining high-accuracy minimum-time trajectories for various Earth-orbit transfers using two different types of low-thrust propulsion models. The first propulsion model assumes the spacecraft has an energy source throughout the entire transfer. This allows for continuous thrust throughout the entire transfer. This type of propulsion model leads to a non-eclipsing orbit transfer. The second propulsion model assumes the spacecraft has an energy source only when it has direct line of sight to the Sun. In other words, the spacecraft will have zero thrust when traveling through the Earth's shadow. This type of propulsion model leads to an eclipsing orbit transfer.

ISBN: 9781369095692Subjects--Topical Terms:

1002622
Aerospace engineering.

Trajectory optimization of low-thrust orbit transfers.
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245 1 0 $a Trajectory optimization of low-thrust orbit transfers.
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300 $a 153 p.
500 $a Source: Dissertation Abstracts International, Volume: 77-12(E), Section: B.
500 $a Adviser: Anil V. Rao.
502 $a Thesis (Ph.D.)--University of Florida, 2015.
520 $a Since the world's first artificial satellite, the Sputnik 1, completed its first orbit around the Earth, we have persistently developed new technology with an emphasis on cost-effective ways to place objects in space. NASA has identified low-thrust propulsion as a critical technology that would greatly reduce the cost of orbital insertion of spacecraft. To date, low-thrust propulsion has been studied in extensive detail and used mostly for interplanetary missions. Little attention, however, has been applied to the usage of low-thrust technology for orbital insertion. This research seeks to devise approaches for obtaining high-accuracy minimum-time trajectories for various Earth-orbit transfers using two different types of low-thrust propulsion models. The first propulsion model assumes the spacecraft has an energy source throughout the entire transfer. This allows for continuous thrust throughout the entire transfer. This type of propulsion model leads to a non-eclipsing orbit transfer. The second propulsion model assumes the spacecraft has an energy source only when it has direct line of sight to the Sun. In other words, the spacecraft will have zero thrust when traveling through the Earth's shadow. This type of propulsion model leads to an eclipsing orbit transfer.
520 $a For the non-eclipsing orbit transfer study, the trajectory optimization problem was posed as a single-phase optimal control problem. An initial guess for the optimal control problem was obtained by solving a sequence of modified optimal control problems where the final true longitude was constrained and the mean square difference between the specified terminal boundary conditions and the computed terminal conditions was minimized. It was found that.
520 $a solutions to the minimum-time low-thrust optimal control problem were only locally optimal in that the solution has essentially the same number of orbital revolutions as that of the initial guess. A search method was then devised that enables computation of solutions with an even lower cost where the final true longitude is constrained to be different from that obtained in the original locally optimal solution. A numerical optimization study was then performed to determine optimal trajectories and control inputs for a range of initial thrust accelerations and constant specific impulses. The key features of the solutions were determined, and relationships were obtained between the optimal transfer time and the optimal final true longitude as a function of the initial thrust acceleration and specific impulse. A detailed post-optimality analysis was performed to verify the close proximity of the numerical solutions to the true optimal solution.
520 $a For the eclipsing orbit transfer study, the trajectory optimization problem was posed as a multiple-phase optimal control problem where the spacecraft can thrust only during phases where it has line of sight to the Sun. Event constraints, based on the geometry of a penumbra shadow region, were enforced between the thrust phases and determine the amount of time spent in an eclipse. An initial guess generation method was developed that constructs an intelligent guess by solving a series of single-phase optimal control problems and analyzing the resulting trajectory to approximate where the spacecraft enters and exits the Earth's shadow. To demonstrate the effectiveness of the approach developed, an optimal transfer trajectory was computed for an Earth-orbit transfer found in the literature. In addition to the comparison case studied, two additional orbit transfers were examined and solutions were presented for various departure dates.
590 $a School code: 0070.
650 4 $a Aerospace engineering. $3 1002622
650 4 $a Mechanics. $3 525881
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