Advanced Space Concepts Laboratory

Glasgow, United Kingdom

Advanced Space Concepts Laboratory

Glasgow, United Kingdom
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Gong S.,Tsinghua University | Li Prof. J.,Tsinghua University | Simo J.,Advanced Space Concepts Laboratory
Journal of Guidance, Control, and Dynamics | Year: 2014

Artificial Lagrange points and periodic orbits around them in the restricted three-body problem have attracted a great deal of attention. Solar sails provide new families of libration points inside regions connected to the classical libration points. Similar to traditional halo orbits centered on the classical libration points, new orbits associated with artificial libration points are widely investigated. The Earth moon libration points have been a topic of great interest in recent years. The orbits around the collinear points are attractive because their unique positions are advantageous for several important applications in space mission design. The out-of-plane distance should be large enough to guarantee that both the lunar far-side and the equatorial regions of the Earth would be visible. Numerical analysis of stability and controllability of the orbits shows that the orbits are unstable but completely controllable with both the reflectivity rate and the sail attitude.


Zamaro M.,Advanced Space Concepts Laboratory | Biggs J.D.,Advanced Space Concepts Laboratory
Proceedings of the International Astronautical Congress, IAC | Year: 2014

One of the paramount stepping stones towards NASA's long-term goal of undertaking human missions to Mars is the exploration of the Martian moons. Since a precursor mission to Phobos would be easier than landing on Mars itself, NASA is targeting this moon for future exploration, and ESA has also announced Phootprint as a candidate Phobos sample-and-return mission. Orbital dynamics around small planetary satellites is particularly complex because many strong perturbations are involved, and the classical circular restricted three-body problem (R3BP) does not provide an accurate approximation to describe the system's dynamics. Phobos is a special case, since the combination of a small mass-ratio and length-scale means that the sphere-of-influence of the moon moves very close to its surface. Thus, an accurate nonlinear model of a spacecraft's motion in the vicinity of this moon must consider the additional perturbations due to the orbital eccentricity and the complete gravity field of Phobos, which is far from a spherical-shaped body, and it is incorporated into an elliptic R3BP using the gravity harmonics series-expansion (ER3BP-GH). In this paper, a showcase of various classes of non-keplerian orbits is identified and a number of potential mission applications in the Mars-Phobos system are proposed: these results could be exploited in upcoming unmanned missions targeting the exploration of this Martian moon. These applications include: low-thrust hovering and orbits around Phobos for close-range observations; the dynamical substitutes of periodic and quasi-periodic Libration Point Orbits in the ER3BP-GH to enable unique low-cost operations for space missions in the proximity of Phobos; their manifold structure for high-performance landing/take-off maneuvers to and from Phobos' surface and for transfers from and to Martian orbits; Quasi-Satellite Orbits for long-period station-keeping and maintenance. In particular, these orbits could exploit Phobos' occulting bulk and shadowing wake as a passive radiation shield during future manned flights to Mars to reduce human exposure to radiation, and the latter orbits can be used as an orbital garage, requiring no orbital maintenance, where a spacecraft could make planned pit-stops during a round-trip mission to Mars. Copyright © 2014 by M. Zamaro and J.D. Biggs.


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.


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.


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.


McKay R.J.,University of Strathclyde | McKay R.J.,Advanced Space Concepts Laboratory | MacDonald M.,University of Strathclyde | MacDonald M.,Advanced Space Concepts Laboratory | And 4 more authors.
Journal of Guidance, Control, and Dynamics | Year: 2011

A survey of highly non-Keplerian orbits with low-thrust propulsion is presented. Keplerian orbits neglect atmospheric drag, solar radiation pressure, nonspherical central bodies, and other perturbations. The term non-Keplerian has been used in reference to orbits where a perturbing or propulsive acceleration acts in addition to that of the effects of gravity. Highly-non-Keplerian orbits can be obtained by considering the dynamics of a low-thrust spacecraft in a rotating frame of reference, where the angular velocity of rotation of the frame of reference is used as a free parameter of the problem. Stationary solutions to the equations of motion can then be sought in this rotating frame of reference, which correspond to periodic, displaced orbits when viewed from an inertial frame of reference. The more thrust that an ion engine can generate, the greater the gravity gradient that can be compensated for and hence the more opportunities there are for applying non-Keplerian orbits.


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.


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.


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

The displaced periodic orbits at a linear order in the circular restricted Earth-Moon system, for which the third massless body uses a hybrid of a solar sail and an SEP system was studied. The first-order approximation is derived for the linearized equations of motion. Then, a feedback linearization control scheme is proposed and implemented. The main idea of this approach is to cancel the nonlinearities and to impose desired linear dynamics satisfied by the solar sail. When the control is applied to the nonlinear system, asymptotic stability is achieved. This provides the key advantage that the displacement distance of the hybrid sail is then constant. A stabilizing approach is then introduced to increase the damping in the system and to allow a higher gain in the controller.


Baig S.,University of Strathclyde | Baig S.,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

This paper discusses a new family of non-Keplerian orbits for solar sail spacecraft displaced above or below the Earth's equatorial plane. The work aims to prove the assertion in the literature that displaced geostationary orbits exist, possibly to increase the number of available slots for geostationary communications satellites. The existence of displaced non-Keplerian periodic orbits is rst shown analytically by linearization of the solar sail dynamics around a geostationary point. The full displaced periodic solution of the nonlinear equations of motion is then obtained using a Hermite-Simpson collocation method with inequality path constraints. The initial guess to the collocation method is given by the linearized solution, and the inequality path constraints are enforced as a box around the linearized solution. The linear and nonlinear displaced periodic orbits are also obtained for the worst-case sun-sail orientation at the solstices. Near-term and high-performance sails can be displaced between 10 and 25 km above the Earth's equatorial plane during the summer solstice, and a perforated sail can be displaced above the usual station-keeping box (75 × 75 km) of nominal geostationary satellites. Light-levitated orbit applications to space solar power are also considered. Copyright © 2009.

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