Molecular characterization and chromosomal localization of two families of satellite DNA in Prochilodus lineatus (Pisces, Prochilodontidae), a species with B chromosomes

Prochilodus lineatus, an abundant species in the Mogi-Guaçu river basin, represents a large part of the region's fishing potential. Karyotypic analyses based on classic cytogenetic techniques have revealed the presence of 54 metasubmetacentric type chromosomes, together with the occurrence of small supernumerary chromosomes with intra and interindividual variations. This paper describes the genomic organization of two families of satellite DNA in the P. lineatus genome. The chromosomal localization these two repetitive DNA families through fluorescence in situ hybridization (FISH) demonstrated that the SATH1 satellite DNA family, composed of approximately 900 bp, was located in the pericentromeric region of a group of chromosomes of the standard complement, as well as on all the B chromosomes. The SATH2 satellite family has a monomeric unit of 441 bp and was located in the pericentromeric regions of some chromosomes of the standard complement, but was absent in the B chromosomes. Double FISH analyses showed that these two families participate jointly in the pericentromeric organization of several chromosomes of this species. The data obtained in this study support the hypothesis that the B chromosomes derive from chromosomes of the standard complement, which are carriers of the SATH1 satellite DNA.

Ten years of NO2 comparisons between ground-based SAOZ and satellite instruments (GOME, SCIAMACHY, OMI)

SAOZ (Systeme d'Analyse par Observations Zenithales) is a ground-based UV-Visible zenith-sky spectrometer installed between 1988 and 1995 at a number of NDSC stations at various latitudes on the globe. The instrument is providing ozone and NO2 vertical columns at sunrise and sunset using the Differential Optical Absorption Spectroscopy (DOAS) technique in the visible spectral range. The ERS-2 GOME Ozone Monitoring Experiment (GOME) in 1995 was the first satellite mission to provide a global picture of atmospheric NO 2 with reasonable spatial and temporal resolution. It was then followed by SCanning ImAging spectroMeter for Atmospheric ChartographY (SCIAMACHY) onboard ENVISAT in 2002, and Ozone Monitoring Instrument (OMI) onboard EOS-AURA in 2004, with a similar capacity to monitor total NO 2. All these instruments are nadir viewing mapping spectrometers, applying the DOAS technique in the visible for deriving the NO2 total column. Here we present the results of NO2 long-term comparisons between GOME and SAOZ for the whole period of GOME operation since 1995 at all latitudes - tropics, mid-latitudes and polar regions - in both hemispheres. Comparisons are also shown with the most recently available SCIAMACHY and OMI data in 2004-2005. Overall, the daytime satellite measurements (around noon) are found consistent with sunrise ground-based data, with an average smaller difference at the tropics and mid-latitudes than in the polar areas in the summer. The agreement is even improved after correcting for the NO2 photochemical change between sunrise and the satellite overpass using a box model. However, some seasonal dependence of the difference between ground-based and satellite total NO2 still remains, related to the accuracy of photochemical simulations and the set of NO2 air mass factors used in the retrievals of both systems.

Real time orbit determination using GPS navigation solution

This work analyses a real time orbit estimator using the raw navigation solution provided by GPS receivers. The estimation algorithm considers a Kalman filter with a rather simple orbit dynamic model and random walk modeling of the receiver clock bias and drift. Using the Topex/Poseidon satellite as test bed, characteristics of model truncation, sampling rates and degradation of the GPS receiver (Selective Availability) were analysed. Copyright © 2007 by ABCM.

On the existence and stability of periodic orbits in non ideal problems: General results

In this work, motivated by non-ideal mechanical systems, we investigate the following O.D.E. ẋ = f (x) + εg (x, t) + ε2g (x, t, ε), where x ∈ Ω ⊂ ℝn, g, g are T periodic functions of t and there is a 0 ∈ Ω such that f (a 0) = 0 and f′ (a0) is a nilpotent matrix. When n = 3 and f (x) = (0, q (x 3) , 0) we get results on existence and stability of periodic orbits. We apply these results in a non ideal mechanical system: the Centrifugal Vibrator. We make a stability analysis of this dynamical system and get a characterization of the Sommerfeld Effect as a bifurcation of periodic orbits. © 2007 Birkhäuser Verlag, Basel.

Study of the potential of irregular shaped bodies and orbits around a non-spherical body

In the present work we show the expressions of the gravitational potential of homogeneous bodies with non-spherical three-dimensional shapes in order to study the trajectories around these bodies. The potentials of a prolate and an oblate ellipsoids with different values of semi-major axis are presented. Their results are validated with a test using a spherical body in order to guarantee the approximation of any body as a polyhedral model of the body. With these expressions we study trajectories of a point of mass around the three-dimensional bodies and the results indicated that there is a group of orbits around those bodies and the polyhedral form of the object does work very well. Copyright IAF/IAA. All rights reserved.

A study of the relative velocities of small particles that are orbiting the earth

The evolution of the velocity of the particles with respect to the circular orbits of satellites that are around the Earth that the particles will cross, suggests a range of possible velocities of impact as a function of the altitude of the satellite. A study made from those results show that the maximum relative velocities occur at the semi-latus rectum, independent of the initial semi-major axis of the particle. Considering both the solar radiation pressure and the oblateness of the Earth, it is visible that a precession in the orbit occurs and there is also a variation in the eccentricity of the particle as a function of its orbital region and its size. This is important information, because the damage caused in a spacecraft depends on the impact velocity.

Third-body perturbation in the case of elliptic orbits for the disturbing body

In the present work it is presented a semi-analytical and a numerical study of the perturbation caused in a spacecraft by a third body using a double averaged analytical model with the disturbing function expanded in Legendre polynomials up to the second-order. The important reason for this procedure is to eliminate the terms due to the short time periodic motion of the spacecraft and to show smooth curves for the evolution of the mean orbital elements for a long time period. The aim of this study is to calculate the effect of lunar perturbations on the orbits of spacecrafts that are traveling around the Earth. It is presented an analysis of the stability of a near-circular orbit and a study to know under which conditions this orbit remains near-circular. A study of the equatorial orbits is also performed.

Influence of nonsphericity of planetary satellites and perturbation of the third-body on the artificial satellites motion

Some orbital characteristics of lunar artificial satellites is presented taking into account the perturbation of the third-body in elliptical orbit and the non-uniform distribution of mass of the Moon. We consider the development of the non-sphericity of the Moon in zonal spherical harmonics up to the ninth order and sectorial harmonic C 22 due to the lunar equatorial ellipticity. The motion of the artificial satellite is studied under the single-averaged analytical model. The average is applied to the mean anomaly of the satellite to analyze low-altitude orbits which are of highest importance for future lunar missions. We found families of frozen orbits with long lifetimes for the problem of an orbiter travelling around the Moon.

Spin stabilized satellite's attitude propagation with quaternions

This paper presents a study of a modeling scheme for the spin stabilized satellites attitude, entirely developed in terms of quaternion parametrization. The analysis includes numerical propagation of the rotational motion equation, considering the influence of the following torques: aerodynamic, gravity gradient, residual magnetic, eddy currents and the one due to the Lorentz force. Applications are developed considering the Brazilian Spin Stabilized Satellites SCD1 and SCD2, which are quite appropriated for verification and comparison of the theory with the real data generated and processed by the INPE's Satellite Control Center (SCC). The results show that for SCD1 and SCD2 the influence of the eddy current torque is bigger than the others ones, not only due to the orbit altitude, but also to other specific satellites characteristics. The influence of the torque due to Lorentz force is smaller than the others ones because of the dimension and the electrical charges of the SCD1 and SCD2. In all performed tests the errors remained within the dispersion range specified for the attitude determination system of INPE's SCC. The results show the feasibility of using the quaternion attitude parametrization for modeling the satellite dynamics of spin stabilized satellites.

Spin stabilized satellite's attitude analytical prediction

An analytical approach for spin stabilized attitude propagation is presented, considering the coupled effect of the aerodynamic torque and the gravity gradient torque. A spherical coordination system fixed in the satellite is used to locate the satellite spin axis in relation to the terrestrial equatorial system. The spin axis direction is specified by its right ascension and the declination angles and the equation of motion are described by these two angles and the magnitude of the spin velocity. An analytical averaging method is applied to obtain the mean torques over an orbital period. To compute the average components of both aerodynamic torque and the gravity gradient torque in the satellite body frame reference system, an average time in the fast varying orbit element, the mean anomaly, is utilized. Afterwards, the inclusion of such torques on the rotational motion differential equations of spin stabilized satellites yields conditions to derive an analytical solution. The pointing deviation evolution, that is, the deviation between the actual spin axis and the computed spin axis, is also availed. In order to validate the analytical approach, the theory developed has been applied for spin stabilized Brazilian satellite SCD1, which are quite appropriated for verification and comparison of the data generated and processed by the Satellite Control Center of the Brazil National Research Institute (INPE). Numerical simulations performed with data of Brazilian Satellite SCD1 show the period that the analytical solution can be used to the attitude propagation, within the dispersion range of the attitude determination system performance of Satellite Control Center of the Brazilian Research Institute.

Optimal low-cost transfer to L4 and L5 lagrangian points based on G-family of periodic orbits

An alternative transfer strategy to send spacecrafts to stable orbits around the Lagrangian equilibrium points L4 and L5 based in trajectories derived from the periodic orbits around LI is presented in this work. The trajectories derived, called Trajectories G, are described and studied in terms of the initial generation requirements and their energy variations relative to the Earth through the passage by the lunar sphere of influence. Missions for insertion of spacecrafts in elliptic orbits around L4 and L5 are analysed considering the Restricted Three-Body Problem Earth- Moon-particle and the results are discussed starting from the thrust, time of flight and energy variation relative to the Earth. Copyright© (2012) by the International Astronautical Federation.

Analytical direct planetary perturbations on the orbital motion of satellites

An analytical expansion of the disturbing function arising from direct planetary perturbations on the motion of satellites is derived. As a Fourier series, it allows the investigation of the secular effects of these direct perturbations, as well as of every argument present in the perturbation. In particular, we construct an analytical model describing the evection resonance between the longitude of pericenter of the satellite orbit and the longitude of a planet, and study briefly its dynamic. The expansion developed in this paper is valid in the case of planar and circular planetary orbits, but not limited in eccentricity or inclination of the satellite orbit. © 2012 Springer Science+Business Media Dordrecht.

Nilpotent orbits and codimension-2 defects of 6d N = (2, 0)theories

We study the local properties of a class of codimension-2 defects of the 6d N = (2, 0) theories of type J = A, D, E labeled by nilpotent orbits of a Lie algebra $g, where g is determined by J and the outer-automorphism twist around the defect. This class is a natural generalization of the defects of the six-dimensional (6d) theory of type SU(N) labeled by a Young diagram with N boxes. For any of these defects, we determine its contribution to the dimension of the Higgs branch, to the Coulomb branch operators and their scaling dimensions, to the four-dimensional (4d) central charges a and c and to the flavor central charge k. © 2013 World Scientific Publishing Company.

Onboard and real-time artificial satellite orbit determination using GPS

An algorithm for real-time and onboard orbit determination applying the Extended Kalman Filter (EKF) method is developed. Aiming at a very simple and still fairly accurate orbit determination, an analysis is performed to ascertain an adequacy of modeling complexity versus accuracy. The minimum set of to-be-estimated states to reach the level of accuracy of tens of meters is found to have at least the position, velocity, and user clock offset components. The dynamical model is assessed through several tests, covering force model, numerical integration scheme and step size, and simplified variational equations. The measurement model includes only relevant effects to the order of meters. The EKF method is chosen to be the simplest real-time estimation algorithm with adequate tuning of its parameters. In the developed procedure, the obtained position and velocity errors along a day vary from 15 to 20 m and from 0.014 to 0.018 m/s, respectively, with standard deviation from 6 to 10 m and from 0.006 to 0.008 m/s, respectively, with the SA either on or off. The results, as well as analysis of the final adopted models used, are presented in this work. © 2013 Ana Paula Marins Chiaradia et al.