IRIS Università degli Studi di Padovahttps://www.research.unipd.itIl sistema di repository digitale IRIS acquisisce, archivia, indicizza, conserva e rende accessibili prodotti digitali della ricerca.Tue, 28 Sep 2021 07:22:34 GMT2021-09-28T07:22:34Z10601Orbits of Selected Globular Clusters in the Galactic Bulgehttp://hdl.handle.net/11577/3287552Titolo: Orbits of Selected Globular Clusters in the Galactic Bulge
Mon, 01 Jan 2018 00:00:00 GMThttp://hdl.handle.net/11577/32875522018-01-01T00:00:00ZThe Equations of Relative Motion in the Orbital Reference Framehttp://hdl.handle.net/11577/2953500Titolo: The Equations of Relative Motion in the Orbital Reference Frame
Abstract: The analysis of relative motion of two spacecraft in Earth-bound orbits is usually carried out on the basis of simplifying assumptions. In particular, the reference spacecraft is assumed to follow a circular orbit, in which case the equations of relative motion are governed by the well-known Hill-Clohessy-Wiltshire (HCW) equations. Circular motion is not, however, a solution when the Earth's flattening is accounted for, except for equatorial orbits, where in any case the acceleration term is not Newtonian. Several attempts have been made to account for the J2 effects, either by ingeniously taking advantage of their differential effects, or by cleverly introducing ad-hoc terms in the equations of motion on the basis of geometrical analysis of the J2 perturbing effects. Analysis of relative motion about an unperturbed elliptical orbit is the next step in complexity. Relative motion about a J2-perturbed elliptic reference trajectory is clearly a challenging problem, which has received little attention. All these problems are based on either the HCW equations for circular reference motion, or the de Vries/Tschauner-Hempel equations for elliptical reference motion, which are both approximate versions of the exact equations of relative motion. The main difference between the exact and approximate forms of these equations consists in the expression for the angular velocity and the angular acceleration of the rotating reference frame with respect to an inertial reference frame. The rotating reference frame is invariably taken as the local orbital frame, i.e., the RTN frame generated by the radial, the transverse, and the normal directions along the primary spacecraft orbit. Some authors have tried to account for the non-constant nature of the angular velocity vector, but have limited their correction to a mean motion value consistent with the J2 perturbation terms. However, the angular velocity vector is also affected in direction, which causes precession of the node and the argument of perigee, i.e., of the entire orbital plane. Here we provide a derivation of the exact equations of relative motion by expressing the angular velocity of the RTN frame in terms of the state vector of the reference spacecraft. As such, these equations are completely general, in the sense that the orbit of the reference spacecraft need only be known through its ephemeris, and therefore subject to any force field whatever. It is also shown that these equations reduce to either the Clohessy-Wiltshire, or the HCW equations, depending on the level of approximation. The explicit form of the equations of relative motion with respect to a J2-perturbed reference orbit is also introduced.
Tue, 01 Jan 2013 00:00:00 GMThttp://hdl.handle.net/11577/29535002013-01-01T00:00:00ZOptimal Maintenance of Relative Circular Inertial Motion for Nulling Interferometry Applicationshttp://hdl.handle.net/11577/2953501Titolo: Optimal Maintenance of Relative Circular Inertial Motion for Nulling Interferometry Applications
Abstract: Spacecraft Formation Flying (FF) is a modern technology in Space Mission Design. Many recent missions have been designed based on this new architecture, in which a number of small, cost-effective satellites cooperate in space in order to achieve certain mission objectives. However, to ensure the feasibility of this paradigm, engineers have to deal with plenty of issues such as the necessity to maintain the formation geometry in a perturbed environment. A solution to this problem is given by optimal control theory, which allows to find the acceleration profile required to reach and keep the desired configuration in the most efficient way (e.g., by minimizing fuel consumption). Accordingly, a Linear Quadratic Tracker (LQT) and a Linear Quadratic Regulator (LQR) has been developed to study the formation keeping of two satellites in a J2-perturbed elliptic orbit. Despite a majority of investigations relies on the inaccurate Hill-Clohessy-Wiltshire model of relative motion, it will be shown that a more realistic dynamical model of the system dynamics leads to appreciable fuel savings. As a test case, a two-satellite configuration has been selected in which one spacecraft---the follower---orbits around the other---the leader---in a plane perpendicular to the line of sight to a chosen star, e.g. $\alpha$Cen, following a circular path covered in a given period. This is akin to the case of nulling interferometry on an Earth-bound orbit, an analogue of the back-up configuration of the L2-positioned Darwin mission. Here we study one of the several spokes or one collector-combiner pair of the full mission configuration. In particular, we first illustrate a Numerical Model (NM) of the relative motion that has been used in the simulations. This model is derived in the Earth Centered Inertial frame (ECI) and is valid for any eccentric orbit. Furthermore, it takes into account the second zonal harmonic of the Earth's gravitational field. Second, the optimal control profile for maintaining the configuration is found by solving the Differential Riccati Equation (DRE) either with the LQT or the LQR. Once the best control scheme is determined and the accelerations found, the $\Delta V$ required per orbit can be estimated and different configurations can be compared. Accordingly, the most efficient formations to satisfy the nulling interferometry requirements for satellites in Earth orbit can be determined. For example, we have investigated the difference in fuel consumption by varying the eccentricity and semimajor axis of the leader spacecraft. In addition, a parametric survey has been performed to quantify the effects of the target choice and the inclination of the leader's orbit on the total formation keeping cost. That is, the Delta_V required per orbit has been calculated as a function of right ascension and declination of the target and for six different inclination values.
Tue, 01 Jan 2013 00:00:00 GMThttp://hdl.handle.net/11577/29535012013-01-01T00:00:00ZGOCE Fully-Dynamic Precise Orbit Recoveryhttp://hdl.handle.net/11577/2953502Titolo: GOCE Fully-Dynamic Precise Orbit Recovery
Abstract: GOCE was launched in 2009 at 250 km altitude to recover Earth's static gravity field. As part of the GOCE-Italy project, we carried out GPS-based, fully-dynamic POD of GOCE for daily arcs covering about 500 days (November 1, 2009 - May 31, 2011). Two sequences were defined and implemented with the software NAPEOS (ESA/ESOC). The first sequence performs the POD task for 30-hours daily arcs, leading to 6 hours overlap between subsequent days, while the second sequence performs the same task for 24-hours daily arcs. Both the sequences were built using the official kinematic Precise Scientific Orbits (PSO) as a-priori orbits. The sequence with overlaps was used to process all the available data, while the second one was run only for those days where the overlaps were not possible or when the first sequence failed. The POD task was successfully performed for 99.4% of the total available days of data, and results show an average post-fit RMS of zero difference phase measurements below 10 mm for 91.5% of the daily arcs. Most orbits compare to less than 6 cm 3D RMS with respect to the official kinematic PSO orbits and the overlapping arcs show an RMS of the distance of about 13 mm.
Tue, 01 Jan 2013 00:00:00 GMThttp://hdl.handle.net/11577/29535022013-01-01T00:00:00ZAnalysis of Steins Cratering History Using the OSIRIS/ROSETTA Imageshttp://hdl.handle.net/11577/2521868Titolo: Analysis of Steins Cratering History Using the OSIRIS/ROSETTA Images
Abstract: We present a preliminary analysis of the craters on the asteroid Steins images obtained by OSIRIS, the imaging system on board the ESA mission ROSETTA, during the flyby on 5th September 2008. Steins has been observed at the closest distance of about 800 km achieving the maximum resolution of 80 m/px. Several small-to-medium craters have been identified, in addition to few very large craters: one of them is nearly 2 km across. The images show also the superposition of small craters on larger ones, and some structure that may represent the remnant of old degraded craters. A structure of chain-like craters has also been identified.
All the craters have been counted in order to get the cumulative number per square km.
Then we have applied our model to estimate the collisional age of Steins using the most recent modeling of the current population of the Main Belt asteroids (Bottke et al., 2005) to define the impactor flux. The model uses the scaling law of Holsapple and Housen (2007) to determine the crater diameter as a function of the impactor radius.
Thu, 01 Jan 2009 00:00:00 GMThttp://hdl.handle.net/11577/25218682009-01-01T00:00:00ZOn the Optimality of a shape-based Approach based on Pseudo-equinoctial elementshttp://hdl.handle.net/11577/2521288Titolo: On the Optimality of a shape-based Approach based on Pseudo-equinoctial elements
Sun, 01 Jan 2006 00:00:00 GMThttp://hdl.handle.net/11577/25212882006-01-01T00:00:00ZOrbit restitution capability of a multiple-antenna GNSS receiver on a highly elliptic orbit reaching above gnss altitudehttp://hdl.handle.net/11577/2521977Titolo: Orbit restitution capability of a multiple-antenna GNSS receiver on a highly elliptic orbit reaching above gnss altitude
Abstract: Astronomical missions are often characterized by high altitude, highly elliptic orbits. We report on the results of a study on the orbit determination capability of a receiver equipped with several GNSS antennas on a 1,000 km by 25,000 km altitude orbit. Detailed visibility analysis shows how this antenna array can help extend the tracking periods to GNSS constellations. Account is taken of the side lobe radio link allowed by the real GPS antennas radiation pattern. High accuracy orbit determination in the few centimeters range is shown to be possible due to the smooth character of the force field, even in the presence of unmodeled attitude variations.
Sun, 01 Jan 2012 00:00:00 GMThttp://hdl.handle.net/11577/25219772012-01-01T00:00:00ZFormation keeping and maneuvering for astronomical, dual spacecraft formation-flying missionshttp://hdl.handle.net/11577/2521997Titolo: Formation keeping and maneuvering for astronomical, dual spacecraft formation-flying missions
Thu, 01 Jan 2009 00:00:00 GMThttp://hdl.handle.net/11577/25219972009-01-01T00:00:00ZComparison among spherical harmonics synthesis methods for functionals of the gravity fieldhttp://hdl.handle.net/11577/2527068Titolo: Comparison among spherical harmonics synthesis methods for functionals of the gravity field
Sat, 01 Jan 2005 00:00:00 GMThttp://hdl.handle.net/11577/25270682005-01-01T00:00:00ZAn Assessment of the Benefits of Including GLONASS Data in GPS-Based Precise Orbit Determination - I: S/A Analysishttp://hdl.handle.net/11577/1468001Titolo: An Assessment of the Benefits of Including GLONASS Data in GPS-Based Precise Orbit Determination - I: S/A Analysis
Sat, 01 Jan 2000 00:00:00 GMThttp://hdl.handle.net/11577/14680012000-01-01T00:00:00Z