Moon-tracking orbits using motorized tethers for continuous earth–moon payload exchanges

For human colonization of the moon to become reality, an efficient and regular means of exchanging resources between the Earth and the moon must be established. One possibility is to pass and receive payloads at regular intervals between a symmetrically laden motorized momentum-exchange tether orbiting about Earth and a second orbiting about the moon. There are significant challenges associated with this method, among the greatest of which is the development of a system that incorporates the complex motion of the moon into its operational architecture in addition to conducting these exchanges on a per-lunar-orbit basis. One way of achieving this is to use a motorized tether orbiting Earth and tracking the nodes of the moon’s orbit to allow payload exchanges to be undertaken periodically with the arrival of the moon at either of these nodes. Tracking these nodes is achieved by arranging the tether to orbit Earth with a critical inclination, thus rendering its argument of perigee stationary in addition to using the precession effects resulting from an oblate Earth. Using this in conjunction with pre-emptive adjustments to its angle of right ascension, the tether will periodically realign itself with these nodes simultaneously with the arrival of the moon

Santa María de Siones

Copia digital. Valladolid : Junta de Castilla y León. Consejería de Cultura y Turismo, 2009-201

Could the Pioneer anomaly have a gravitational origin?

If the Pioneer anomaly has a gravitational origin, it would, according to the equivalence principle, distort the motions of the planets in the Solar System. Since no anomalous motion of the planets has been detected, it is generally believed that the Pioneer anomaly can not originate from a gravitational source in the Solar System. However, this conclusion becomes less obvious when considering models that either imply modifications to gravity at long range or gravitational sources localized to the outer Solar System, given the uncertainty in the orbital parameters of the outer planets. Following the general assumption that the Pioneer spacecraft move geodesically in a spherically symmetric spacetime metric, we derive the metric disturbance that is needed in order to account for the Pioneer anomaly. We then analyze the residual effects on the astronomical observables of the three outer planets that would arise from this metric disturbance, given an arbitrary metric theory of gravity. Providing a method for comparing the computed residuals with actual residuals, our results imply that the presence of a perturbation to the gravitational field necessary to induce the Pioneer anomaly is in conflict with available data for the planets Uranus and Pluto, but not for Neptune. We therefore conclude that the motion of the Pioneer spacecraft must be non-geodesic. Since our results are model independent within the class of metric theories of gravity, they can be applied to rule out any model of the Pioneer anomaly that implies that the Pioneer spacecraft move geodesically in a perturbed spacetime metric, regardless of the origin of this metric disturbance.Comment: 16 pages, 6 figures. Rev. 3: Major revision. Accepted for publication in Phys. Rev. D. Rev. 4: Added two reference

All-propulsion design of the drag-free and attitude control of the European satellite GOCE

This paper concerns the drag-free and attitude control (DFAC) of the European Gravity field and steady-state Ocean Circulation Explorer satellite (GOCE), during the science phase. GOCE aims to determine the Earth's gravity field with high accuracy and spatial resolution, through complementary space techniques such as gravity gradiometry and precise orbit determination. Both techniques rely on accurate attitude and drag-free control, especially in the gradiometer measurement bandwidth (5-100mHz), where non-gravitational forces must be counteracted down to micronewton, and spacecraft attitude must track the local orbital reference frame with micro-radian accuracy. DFAC aims to enable the gravity gradiometer to operate so as to determine the Earth's gravity field especially in the so-called measurement bandwidth (5-100mHz), making use of ion and micro-thruster actuators. The DFAC unit has been designed entirely on a simplified discrete-time model (Embedded Model) derived from the fine dynamics of the spacecraft and its environment; the relevant control algorithms are implemented and tuned around the Embedded Model, which is the core of the control unit. The DFAC has been tested against uncertainties in spacecraft and environment and its code has been the preliminary model for final code development. The DFAC assumes an all-propulsion command authority, partly abandoned by the actual GOCE control system because of electric micro-propulsion not being fully developed. Since all-propulsion authority is expected to be imperative for future scientific and observation missions, design and simulated results are believed to be of interest to the space communit

Orbital Debris-Debris Collision Avoidance

We focus on preventing collisions between debris and debris, for which there is no current, effective mitigation strategy. We investigate the feasibility of using a medium-powered (5 kW) ground-based laser combined with a ground-based telescope to prevent collisions between debris objects in low-Earth orbit (LEO). The scheme utilizes photon pressure alone as a means to perturb the orbit of a debris object. Applied over multiple engagements, this alters the debris orbit sufficiently to reduce the risk of an upcoming conjunction. We employ standard assumptions for atmospheric conditions and the resulting beam propagation. Using case studies designed to represent the properties (e.g. area and mass) of the current debris population, we show that one could significantly reduce the risk of nearly half of all catastrophic collisions involving debris using only one such laser/telescope facility. We speculate on whether this could mitigate the debris fragmentation rate such that it falls below the natural debris re-entry rate due to atmospheric drag, and thus whether continuous long-term operation could entirely mitigate the Kessler syndrome in LEO, without need for relatively expensive active debris removal.Comment: 13 pages, 8 figures. Accepted for publication in Advances in Space Researc

Renán Arcadio Poveda Ricalde

Renán Arcadio Poveda Ricalde nace en la ciudad de Mérida Yucatán el 15 de julio de 1930. Durante su niñez, acostumbraba caminar, del lado de su padre, viendo los atardeceres en las playas del puerto de Progreso. Al caer la noche, disfrutaba el maravilloso espectáculo de la bóveda celeste, e incansablemente lo cuestionaba sobre los nombres de las estrellas, de que estaban hechas y que tan lejos se encontraban. Esas caminatas lo marcarían por el resto de su vida y su curiosidad por el universo se transformaría en su pasión

Variation of Area-to-Mass-Ratio of HAMR Space Debris Objects

An unexpected space debris population has been detected in 2004 Schildknecht et al. (2003, 2004) with the unique properties of a very high area-to-mass ratio (HAMR) Schildknecht et al. (2005a). Ever since it has been tried to investigate the dynamical properties of those objects further. The orbits of those objects are heavily perturbed by the effect of direct radiation pressure. Unknown attitude motion complicates orbit prediction. The area-to-mass ratio of the objects seems to be not stable over time. Only sparse optical data is available for those objects in drift orbits. The current work uses optical observations of five HAMR objects, observed over several years and investigates the variation of their area-to-mass ratio and orbital parameters. A normalized orbit determination setup has been established and validated with two low and two of the high ratio objects, to ensure, that comparable orbits over longer time spans are determined even with sparse optical data.Comment: 10 pages, accepted Monthly Notices of Royal Astronomical Society, MN-11-1785-MJ.R1, The definitive version is available at www.blackwell-synergy.co

Analytical sun synchronous low-thrust manoeuvres

Article describes analytical sun synchronous low-thrust manoeuvres

Extension of the sun-synchronous Orbit

Through careful consideration of the orbit perturbation force due to the oblate nature of the primary body a secular variation of the ascending node angle of a near-polar orbit can be induced without expulsion of propellant. Resultantly, the orbit perturbations can be used to maintain the orbit plane in, for example, a near-perpendicular (or at any other angle) alignment to the Sun-line throughout the full year of the primary body; such orbits are normally termed Sun-synchronous orbits [1, 2]. Sun-synchronous orbits about the Earth are typically near-circular Low-Earth Orbits (LEOs), with an altitude of less than 1500 km. It is normal to design a LEO such that the orbit period is synchronised with the rotation of the Earth‟s surface over a given period, such that a repeating ground-track is established. A repeating ground-track, together with the near-constant illumination conditions of the ground-track when observed from a Sun-synchronous orbit, enables repeat observations of a target over an extended period under similar illumination conditions [1, 2]. For this reason, Sun-synchronous orbits are extensively used by Earth Observation (EO) platforms, including currently the Environmental Satellite (ENVISAT), the second European Remote Sensing satellite (ERS-2) and many more. By definition, a given Sun-synchronous orbit is a finite resource similar to a geostationary orbit. A typical characterising parameter of a Sun-synchronous orbit is the Mean Local Solar Time (MLST) at descending node, with a value of 1030 hours typical. Note that ERS-1 and ERS-2 used a MLST at descending node of 1030 hours ± 5 minutes, while ENVISAT uses a 1000 hours ± 5 minutes MLST at descending node [3]. Following selection of the MLST at descending node and for a given desired repeat ground-track, the orbit period and hence the semi-major axis are fixed, thereafter assuming a circular orbit is desired it is found that only a single orbit inclination will enable a Sun-synchronous orbit [2]. As such, only a few spacecraft can populate a given repeat ground-track Sun-synchronous orbit without compromise, for example on the MLST at descending node. Indeed a notable feature of on-going studies by the ENVISAT Post launch Support Office is the desire to ensure sufficient propellant remains at end-of-mission for re-orbiting to a graveyard orbit to ensure the orbital slot is available for future missions [4]. An extension to the Sun-synchronous orbit is considered using an undefined, non-orientation constrained, low-thrust propulsion system. Initially the low-thrust propulsion system will be considered for the free selection of orbit inclination and altitude while maintaining the Sun-synchronous condition. Subsequently the maintenance of a given Sun-synchronous repeat-ground track will be considered, using the low-thrust propulsion system to enable the free selection of orbit altitude. An analytical expression will be developed to describe these extensions prior to then validating the analytical expressions within a numerical simulation of a spacecraft orbit. Finally, an analysis will be presented on transfer and injection trajectories to these orbits

Interplanetary program to optimize simulated trajectories (IPOST). Volume 4: Sample cases

The Interplanetary Program to Optimize Simulated Trajectories (IPOST) is intended to support many analysis phases, from early interplanetary feasibility studies through spacecraft development and operations. The IPOST output provides information for sizing and understanding mission impacts related to propulsion, guidance, communications, sensor/actuators, payload, and other dynamic and geometric environments. IPOST models three degree of freedom trajectory events, such as launch/ascent, orbital coast, propulsive maneuvering (impulsive and finite burn), gravity assist, and atmospheric entry. Trajectory propagation is performed using a choice of Cowell, Encke, Multiconic, Onestep, or Conic methods. The user identifies a desired sequence of trajectory events, and selects which parameters are independent (controls) and dependent (targets), as well as other constraints and the cost function. Targeting and optimization are performed using the Standard NPSOL algorithm. The IPOST structure allows sub-problems within a master optimization problem to aid in the general constrained parameter optimization solution. An alternate optimization method uses implicit simulation and collocation techniques