One promising option for space operational responsiveness is orbital maneuvering. In low orbit, a maneuverable spacecraft can provide valuable benefits such as better coverage properties, increased revisit times, selectable targets, and local over-flight times. Such maneuvers are not common due to the high cost of chemical propulsion. The more recent paradigm of Responsive Space is to rapidly launch a small, inexpensive asset and use it in a short, disposable fashion. This concept relies on drastically reducing the cost of launch, yet it remains the most expensive piece, so additional cost savings can be realized by minimizing the need for launches. Electric propulsion has been considered as an efficient alternative to chemical propulsion. With technological advances, electric propulsion can provide responsiveness in a timely, fuel-efficient manner without requiring repeated launches to satisfy multiple missions. This study shows the control algorithm for a single low-Earth satellite equipped with the proper electric propulsion to overfly any target inside its coverage area in as little as 34 hours for 1.8 percent of its propellant budget. A comprehensive survey to quantify global reach requirements is provided and the optimal time and fuel solutions are explored. The results strongly support that electric propulsion could be a key enabler in responsive operations. Nomenclature A = perturbing acceleration magnitude, km/s 2 a = semi-major axis, km a h = normal acceleration component with respect to orbital plane, km/s 2 a r = radial acceleration component in local-vertical, local-horizontal frame, km/s 2 a θ = acceleration component completing the right-handed coordinate system, km/s 2 e = eccentricity h = angular momentum ĥ = normal component wrt orbital plane of the local-vertical, local-horizontal frame i = inclination, degrees M = mean anomaly, rad m 0 = initial spacecraft mass, kg m = mass flow rate, kg/s Δm = change in mass due to maneuvering, kg n = mean motion, rad/s p = semi-latus rectum, km r = radius vector, components in km r = magnitude of radius vector, km rˆ = radial component of the local-vertical, local-horizontal frame T = thrust magnitude, N t = time, s t 0 = initial time, s t f = final time, s Δt = change in over-flight time, s u = argument of latitude, degrees V = velocity vector, components in km/s ΔV = velocity change, km/s x = state vector in terms of Classical Orbital Elements x 0 = initial state vector in Classical Orbital Elements γ = angle between Earth-Centered Inertial (ECI) and ECF frames, rad γ g = Greenwich sidereal time, rad θ = out-of-plane thrust angle wrt rˆ-ˆ plane, degrees ˆ = tangential component of the local-vertical, local-horizontal frame λ = latitude, degrees Downloaded by UNIVERSITY OF NEW SOUTH WALES on August 23, 2015 | http://arc.aiaa.org | = Earth's gravitational parameter, 3.98601 x10 5 km 3 /s 2 ν = true anomaly, rad φ = longitude, degrees ψ = thrust control vector, components in degrees ψ = thrust angle in rˆ-ˆ plane, degrees Ω = right ascension of the a...
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