Azimuthal repositioning of payloads in heliocentric orbit using solar sails

FUTURE solar physics missions will require the ability to reposition multiple spacecraft at different azimuthal positions relative to the Earth, while remaining close to a one year circular orbit. Such azimuthal repositioningwill allow stereoscopicviews of solar features to be generated and will allow imaging of coronal mass ejections as they transit the sun-Earth line. The NASA STEREO mission, which is scheduled for launch in 2005, will utilize two spacecraft to perform such tasks. Both spacecraft will be launched on a Delta II 7925 and will use multiple lunar gravity assists to maneuver the spacecraft onto leading and trailing heliocentric orbits. The two spacecraft will then drift ahead of and behind the Earth on free-drift trajectories,with increasingEarth-sun-spacecraft angles

A continuum model for the orbit evolution of self-propelled 'smart dust' swarms

A continuity equation is developed to model the evolution of a swarm of self-propelled ‘smart dust’ devices in heliocentric orbit driven by solar radiation pressure. These devices are assumed to be MEMs-scale (micro-electromechanical systems) with a large area-to-mass ratio. For large numbers of devices it will be assumed that a continuum approximation can be used to model their orbit evolution. The families of closed-form solutions to the resulting swarm continuity equation then represent the evolution of the number density of devices as a function of both position and time from a set of initial data. Forcing terms are also considered which model swarm sources and sinks (device deposition and device failure). The closed-form solutions presented for the swarm number density provide insights into the behaviour of swarms of self-propelled ‘smart dust’ devices an can form the basis of more complex mission design methodologies

On-orbit assembly using superquadric potential fields

The autonomous on-orbit assembly of a large space structure is presented using a method based on superquadric artificial potential fields. The final configuration of the elements which form the structure is represented as the minimum of some attractive potential field. Each element of the structure is then considered as presenting an obstacle to the others using a superquadric potential field attached to the body axes of the element. A controller is developed which ensures that the global potential field decreases monotonically during the assembly process. An error quaternion representation is used to define both the attractive and superquadric obstacle potentials allowing the final configuration of the elements to be defined through both relative position and orientation. Through the use of superquadric potentials, a wide range of geometric objects can be represented using a common formalism, while collision avoidance can make use of both translational and rotation maneuvers to reduce total maneuver cost for the assembly process

Agile solar sailing in three-body problem : Motion between artificial equilibrium points

This paper proposes a range of time-optimal, solar sail trajectories between artificial equilibria in the Sun-Earth three body system to create an agile solar sailing mission. This allows different mission objectives to be fulfilled at different AEPs during different stages of the mission. The analyses start from a solar sail at the sub-L1 point (sunward of the classical L1 point) which is targeted by NASA’s Sunjammer mission (launch in 2014) for advanced space weather warning. From this sub-L1 point, trajectories are investigated that: 1) take the solar sail to an AEP in the ecliptic plane, but slightly trailing the Earth to be ahead of the Earth in the Parker spiral to potentially increase space weather warning times even further; 2) take the solar sail to and between AEPs displaced above or below the ecliptic plane for high-latitude observations; 3) take the solar sail from the vicinity of the L1 point to the vicinity of the L¬2 point for additional Earth observations, geomagnetic tail investigations and astronomical observations. To find time-optimal trajectories, the optimal control problem associated with each of the transfers is defined and solved using a direct pseudospectral method. The resulting time of flights are reasonable, ranging from 85 days to 232 days, and the transfers are very smooth, requiring only a minimum solar sail steering effort in most cases. Since all results are generated for a sail performance equal to that of the Sunjammer sail, the proposed trajectories provide interesting end-of-mission opportunities for the Sunjammer sail after it retires at the sub-L1 point

Reconfiguration of a four-bar mechanism using phase space connections

Linkage mechanisms are perhaps the simplest mechanical structures in engineering, but they can exhibit significant nonlinearity which can in principle be exploited. In this paper a simple smart structure model is developed based on such nonlinearity to investigate the reconfiguration of a four-bar mechanism through phase space connections. The central idea is based on heteroclinic connections in the mechanism phase space between equal-energy unstable equilibria. It is proposed that transitions between such equal-energy unstable (but actively controlled) equilibria in principle require zero net energy input, compared to transitions between stable equilibria which require the input and then dissipation of energy. However, it can be difficult to obtain such heteroclinic connections numerically in complex dynamical systems, therefore an objective function approach is used to seek transtions between unstable equilibria which approximate true heteroclinic connections. The instability inherent in the model is therefore actively utilised to provide energy-efficient transitions between configurations of the mechanism. It will be shown that the four-bar mechanism then forms the basis for an elastic model of a smart buckling beam

Space-Enhanced Solar Power for Equatorial Regions

This paper examines the concept of solar mirrors in a Earth orbit to provide solar farms with additional solar power during the hours of darkness. The design of the orbit is key for the purposes of the mission: the mirror needs continuous access to the Sun and the solar farm simultaneously. Therefore, orbits with high-eccentricity will be considered to increase the visibility time. Also, since the most convenient locations for solar power farms are about the equator, a suitable orbit should have a low inclination. This issue can be addressed through the concept of anti-heliotropic orbits that exploits mainly solar radiation pressure perturbations to generate highly-eccentric equatorial orbits able to maintain the orientation with respect to the Sun. The considered configuration consists in two space mirrors in a flower constellation rotating with the Earth to deliver a repeat ground track

Artificial three-body equilibria for hybrid low-thrust propulsion

This paper proposes a new concept of creating artificial equilibrium points in the circular restricted three body problem, where the third body uses a hybrid of solar sail and solar electric propulsion. The work aims to investigate the use of a hybrid sail for artificial equilibrium points that are technologically di±cult with either of these propulsion systems alone. The hybrid sail has freedom in specifying the sail lightness number, then minimizing the required thrust acceleration from the solar electric propulsion thruster while satisfying the equilibrium condition. The stability analysis of such artificial equilibrium points by a linear method results in a linear time varying (mass) system. The freezing time method then provides unstable and marginally stable regions for hybrid solar sail artificial equilibria. We compare these propulsion systems with a given payload mass and mission life for a polar observation mission. For a near term sail assembly loading we find for the hybrid sail a substantially lower propellant mass compared to solar electric propulsion and lower sail length with respect to a solar sail, and a lower initial spacecraft mass

Spacecraft formation-flying using potential functions

A group of small spacecraft able to change their relative position and attitude through the use of the potential function method is discussed. The spacecraft shapes, sizes and manoeuvring capabilities are not identical, although all are assumed to manoeuvre using continuous thrusters. A hyperbolic form of the attractive potential function is used to reduce actuator effort by using natural orbital motion to approaching the goal configuration. A superquadric repulsive potential with 3D a rigid object representation is then used to provide an accurate representation of the shape of spacecraft in the potential function. As the spacecraft start away from their goal, a parabolic attractive potential is inefficient as the control force increases with distance from the goal. Using a hyperbolic attractive potential, the control force is independent of the distance to goal, ensuring smooth manoeuvring towards the goal with a bound actuator effort

Displaced solar sail orbits : dynamics and applications

We consider displaced periodic orbits at linear order in the circular restricted Earth-Moon system, where the third massless body is a solar sail. These highly non-Keplerian orbits are achieved using an extremely small sail acceleration. Prior results have been developed by using an optimal choice of the sail pitch angle, which maximises the out-of-plane displacement. In this paper we will use solar sail propulsion to provide station-keeping at periodic orbits around the libration points using small variations in the sail's orientation. By introducing a first-order approximation, periodic orbits are derived analytically at linear order. These approximate analytical solutions are utilized in a numerical search to determine displaced periodic orbits in the full nonlinear model. Applications include continuous line-of-sight communications with the lunar poles

Design of ballistic three-body trajectories for continuous polar earth observation in the earth-moon system

This paper investigates orbits and transfer trajectories for continuous polar Earth observation in the Earth-Moon system. The motivation behind this work is to complement the services offered by polar-orbiting spacecraft, which offer high resolution imaging but poor temporal resolution, due to the fact that they can only capture one narrow swath at each polar passage. Conversely, a platform for high-temporal resolution imaging can enable a number of applications, from accurate polar weather forecasting to Aurora study, as well as direct-link telecommunications with high-latitude regions. Such a platform would complement polar orbiters. In this work, we make use of resonant gravity swing-by manoeuvres at the Moon in order to design trajectories that are suitable for quasi-continuous polar observation. In particular, it is shown that the Moon can flip the line of apsides of a highly eccentric, highly inclined orbit from north to south, without the need for thrust. In this way, a spacecraft can alternatively loiter for an extended period of time above the two poles. In addition, at the lunar encounter it is possible to change the period of time spent on each pole. In addition, we also show that the lunar swing-by can be exploited for transfer to a so-called pole-sitter orbit, i.e. a spacecraft that constantly hovers above one of the Earth's poles using continuous thrust. It is shown that, by using the Moon's gravity to change the inclination of the transfer trajectory, the total Δv is less than using a trajectory solely relying on high-thrust or low-thrust, therefore enabling the launchers to inject more mass into the target pole-sitter position