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Port Glasgow, United Kingdom

McInnes C.R.,University of Strathclyde | McInnes C.R.,Advanced Space Concepts Laboratory
Journal of Guidance, Control, and Dynamics | Year: 2014

An approximate closed-form solution is presented for solar sail spiral trajectories with sail degradation. Because exposure to the space environment is cumulative, the impact of degradation on the sail thrust magnitude forms an integral function over the spiral duration. The time evolution of the solar sail spiral trajectory is, therefore, described by an integro-differential equation that, for this problem, does not apparently possess an explicit closed-form solution. The optimum fixed sail pitch angle is now a trade-off between maximizing the transverse component of sail thrust while minimizing the projected sail area exposed to the flux of solar radiation. It can be seen that the closed-form solutions obtained using the quasi-circular orbit approximation provide an accurate representation of the dynamics of the problem while clearly missing the orbit eccentricity forcing seen in the numerical integration, as is to be expected. The orbit evolution of a solar sail subject to optical degradation has been investigated and an approximate closed-form analytical solution found, assuming a quasi-circular spiral with a fixed sail pitch angle. This new solar sail spiral problem poses challenges because the sail degradation is cumulative, resulting in an integro-differential equation. Source


Sanchez J.P.,University of Strathclyde | McInnes C.,University of Strathclyde | McInnes C.,Advanced Space Concepts Laboratory
Journal of Spacecraft and Rockets | Year: 2011

Most future concepts for the exploration and exploitation of space require a large initial mass in low Earth orbit. Delivering this required mass from the Earth's surface increases cost duetothe large energy input necessary tomove mass out of the Earth's gravity well. An alternative is to search for resources in-situ among the near-Earth asteroid population. The near-Earth asteroid resources that could be transferred to a bound Earth orbit are determined by integrating the probability of finding asteroids inside the Keplerian orbital element space of the set of transfers with an specific energy smaller than a given threshold. Transfers are defined by a series of impulsive maneuvers and computed using the patched-conic approximation. The results show that even moderately-low-energy transfers enable access to a large mass of resources. Source


Lucking C.,University of Strathclyde | Colombo C.,University of Southampton | McInnes C.R.,Advanced Space Concepts Laboratory
Journal of Spacecraft and Rockets | Year: 2013

A deorbiting strategy for small satellites is proposed that exploits the effect of solar radiation pressure to increase the spacecraft orbit eccentricity so that the perigee falls below an altitude where atmospheric drag will cause the spacecraft orbit to naturally decay. This is achieved by fitting the spacecraft with an inflatable reflective balloon. Once this is fully deployed, the overall area-to-mass ratio of the spacecraft is increased; hence, solar radiation pressure and aerodynamic drag have a greatly increased effect on the spacecraft orbit. An analytical model of the orbit evolution due to solar radiation pressure and the J2 effect as a Hamiltonian system show the evolution of an initially circular orbit. The maximum reachable orbit eccentricity as a function of semimajor axis and area-to-mass ratio is found analytically for deorbiting from circular equatorial orbits of different altitudes. The analytical planar model is then adapted for sun-synchronous orbits. The model is validated numerically and verified for three test cases using a highaccuracy orbit propagator. The regions of orbits for which solar radiation pressure-augmented deorbiting is most effective are identified. Finally, different options for the design of the deorbiting device are discussed. Copyright © 2012 by Charlotte Lücking. Published by the American Institute of Aeronautics and Astronautics, Inc. Source


Biggs J.D.,University of Strathclyde | Biggs J.D.,Advanced Space Concepts Laboratory | McInnes C.R.,University of Strathclyde | McInnes C.R.,Advanced Space Concepts Laboratory
Journal of Guidance, Control, and Dynamics | Year: 2010

A study was conducted to demonstrate the use of sail propulsion to stabilize a spacecraft about an artificial libration point. It was demonstrated that the constant acceleration from a solar sail had the potential to be used to generate artificial libration points in the Earth-sun three-body problem. This was achieved by directing the thrust due to the sail such that it added to the centripetal and gravitational forces. A potential alternative to stabilizing a sail about an artificial libration point was to fix the attitude of the solar sail and actuate the lightness number. Sail lightness number β was the ratio of the solar radiation pressure acceleration to the solar gravitational acceleration. This lightness number control, called β, offered the potential of passive stabilization of a sail about artificial libration points. The setting of the circular restricted three-body problem (CRTBP) was also used to assess the use of β control where the massless body was a solar sail and the primaries were the Sun and Earth. Source


Ceriotti M.,University of Strathclyde | Ceriotti M.,Advanced Space Concepts Laboratory | McInnes C.R.,University of Strathclyde | McInnes C.R.,Advanced Space Concepts Laboratory
Journal of Guidance, Control, and Dynamics | Year: 2011

A pole-sitter orbit is a closed path that is constantly above one of the Earth's poles by means of continuous low thrust. This work proposes to hybridize solar sail propulsion and solar electric propulsion on the same spacecraft to enable such a pole-sitter orbit. Locally optimal control laws are found with a semianalytical inverse method, starting from a trajectory that satisfies the pole-sitter condition in the sun-Earth circular restricted three-body problem. These solutions are subsequently used as a first guess to find optimal orbits, using a direct method based on pseudospectral transcription. The orbital dynamics of both the pure solar electric propulsion case and the hybrid case are investigated and compared. It is found that the hybrid spacecraft allows savings on propellant mass fraction. Finally, is it shown that for sufficiently long missions, a hybrid pole sitter, based on midterm technology, enables a consistent reduction in the launch mass for a given payload, with respect to a pure solar electric propulsion spacecraft. Copyright © 2011 by Matteo Ceriotti and Colin R. McInnes. Source

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