Optimization of Multiple-Impulse, Multiple-Revolution Rendezvous Problem Using Social Cognition Optimization

It is generally agreed that there is no direct evidence about the cognitive abilities, linguistic and non-linguistic, of early humans. There is an approach, though, that has been used to fruitfully investigate the linguistic abilities of these humans. Known as the “Windows Approach,” it proceeds from the assumption that inferences about early forms of human language, including their origin and evolution, can be drawn from other phenomena about which there is direct evidence. These other phenomena are taken to be or provide windows on such forms of language. The central idea expounded in the present chapter is that the conceptual constructs of the Windows Approach can be used to investigate non-linguistic abilities of early humans as well. The chapter does this in three steps: first, by setting out in general terms what a window of the kind concerned is; second, by illustrating the use of core concepts, distinctions and principles of the Windows Approach to make and appraise inferences about linguistic abilities of Neandertals; third, by showing how some of these conceptual constructs have been used in accounts of non-linguistic abilities of early humans. The focus here is on work that draws inferences about cognitive abilities of these humans from hafted spears, and from traps and snares used for capturing smaller animals.

Modeling approaches for precise relativistic orbits: Analytical, Lie-series, and pN approximation

Impact of orbit modeling on DORIS station position and Earth rotation estimates

PurposeThe purpose of this paper is to assess the orbital perturbation caused by the gravitational orbit–attitude coupling of spacecraft in the proximity of asteroids.Design/methodology/approachThe gravitational orbit–attitude coupling perturbation (GOACP), which has been neglected before in the close-proximity orbital dynamics about asteroids, is investigated and compared with other orbital perturbations. The GOACP has its origin in the fact that the gravity acting on a non-spherical extended body is actually different from that acting on a point mass located at the body’s center of mass, which is the approximated model in the orbital dynamics. Besides, a case study of a tethered satellite system is given by numerical simulations.FindingsIt is found that the ratio of GOACP to the asteroid’s non-spherical gravity is the order of ρ/ae, where ρ is the spacecraft’s characteristic dimension and ae is the asteroid’s mean radius. It can also be seen that as ρ increases, GOACP will also increase but the solar radiation pressure (SRP) will decrease due to the decreasing area-to-mass ratio. The GOACP will be more significant than SRP at small orbital radii for a large-sized spacecraft. Based on the results by analyses and simulations, it can be concluded that GOACP needs to be considered in the orbital dynamics for a large-sized spacecraft in the proximity of a small asteroid.Practical implicationsThis study is of great importance for the future asteroids missions for scientific explorations and near-Earth objects mitigation.Originality/valueThe GOACP, which has been neglected before, is revealed and studied.

The electrostatic accelerometer, assembled on gravity satellite, serves to measure all non-gravitational accelerations caused by atmosphere drag or solar radiation pressure, etc. The proof-mass center of the accelerometer needs to be precisely positioned at the center of gravity satellite, otherwise, the offset between them will bring measurement disturbance due to angular acceleration of satellite and gradient. Because of installation and measurement errors on the ground, fuel consumption during the in-flight phase and other adverse factors, the offset between the proof-mass center and the satellite center of mass is usually large enough to affect the measurement accuracy of the accelerometer, even beyond its range. Therefore, the offset needs to be measured or estimated, and then be controlled within the measurement requirement of the accelerometer by the center of mass (COM) adjustment mechanism during the life of the satellite. The estimation algorithm based on EKF, which uses the measurement of accelerometer, gyro and magnetometer, is put forward to estimate the offset, and the COM adjustment mechanism then adjusts the satellite center of mass in order to make the offset meet the requirement. With the special configuration layout, the COM adjustment mechanism driven by the stepper motors can separately regulate X, Y and Z axes. The associated simulation shows that the offset can be controlled better than 0.03 mm for all the axes with the method mentioned above.

Solar radiation pressure on (99942) Apophis

This paper presents the results of a numerical evaluation of the natural lifetime reduction in low Earth orbit, due to dynamical perturbations. The study considers two values for the area-to-mass ratio, a nominal ratio which resembles a typical value of spacecraft in orbit today, and an enhanced ratio which covers the surface augmentation. The results were obtained with two orbit propagators, one of a semi-analytical nature and the second one using non-averaged equations of motion. The simulations for both propagators were set up similarly to allow comparison. They both use the solar radiation pressure and the secular terms of the geopotential (J2,J4 and J6). The atmospheric drag was turned on and off in both propagators to alternatively study the eccentricity build up and the residual lifetime. The non-averaging case also covers a validation with the full 6 × 6 geopotential. The results confirm the findings in previous publications, that is, the possibility for de-orbiting from altitudes above the residual atmosphere if a solar sail is deployed at the end-of-life, due to the combined effect of solar radiation pressure and the oblateness of the Earth. At near polar inclinations, shadowing effects can be exploited to the same end. The results obtained with the full, non-averaging propagator revealed additional de-orbiting corridors associated with solar radiation pressure which were not found by previous work on space debris mitigation. The results of both tools are compared for specific initial conditions. For nominal values of area-to-mass ratio, instead, it is confirmed that this resonance effect is negligible.The paper then puts the findings in the perspective of the current satellite catalogue. It identifies space missions which are currently close to a resonance corridor and shows the orbit evolution within the resonances with a significantly shorter residual orbital lifetime. The paper finishes with a discussion on the exploitation of these effects with regards to the long-term simulation of the space debris environment and a flux and collision probability comparison.

Dynamical models and the onset of chaos in space debris

Since the beginning of satellite altimetry missions, the ocean surface topography community requires precise and accurate satellite orbits. With start of the satellite Sentinel-3A on February 16, 2016, two radio- and one altimeter onboard the satellite accomplish the Copernicus program with a ocean- and land monitoring mission, planned for a nominal mission lifetime of 7 years. The satellite is orbiting the Earth on a polar, Sun-synchronous trajectory at an altitude of 815 km. For the purpose of precise orbit determination, the satellite is equipped with a geodetic-grade dual-frequency Global Positioning System (GPS) receiver. The GPS measurements are employed together with a set of gravitational and non-gravitational models in a Reduced-Dynamic Orbit Determination (RDOD) approach, which combines the advantages of a dynamic and a kinematic positioning for deriving precise satellite orbits. However, especially the non-gravitational force models require sophisticated modeling techniques. Therefore, a satellite macro model is introduced, which allows a proper modeling of accelerations due to Solar Radiation Pressure (SRP), Earth Radiation Pressure (ERP), and atmospheric drag. Especially the Sun-synchronous orbit, and the huge solar array, which is loosely coupled to the satellite body, makes the precise orbit determination challenging. As members of the Copernicus Quality Working Group, the Astronomical Institute of the University of Bern (AIUB), and the German Aerospace Center (DLR) are, among others, responsible for the orbit validation of Sentinel-3A. Both groups make use of a satellite macro model within a reduced-dynamic approach but differ in the employed software solutions and the pseudo-stochastic modeling. Basically, the pseudo-stochastic parameters allow to compensate potential deficits in the employed force models. Within this poster, the modeling aspects are briefly introduced, followed by comparing the results of both groups. The results include the estimated empirical accelerations, and the estimated satellite orbits of Sentinel-3A. Furthermore, the orbit quality is assessed by Satellite Laser Ranging (SLR), an external and independent tool for orbit validation.

This paper studies the orbital perturbation effects on the azimuthal resolution in lunar-based synthetic aperture radar (LBSAR). We derive explicit expressions for the Doppler frequency modulation rate (DFMR) and beam-crossing velocity using the antenna beam pointing and orbit models. Following that, the azimuthal resolution is expressed in line with orbital elements and SAR configurations. The results show that the long-term orbital variations caused by accumulated perturbation effects significantly affect the azimuthal resolution, which, in effect, produces aperiodic variations in the azimuthal resolution. Such a phenomenon is most distinguished for a large LBSAR look angle, leading to a fluctuation of over 30% or even larger in the azimuthal resolution across different cycles. Additionally, the errors give rise by short-term orbital perturbations could impact azimuthal resolution to a lesser extent, with corresponding fluctuations consistently below 3%. The findings reveal that it is imperative to consider the irregular variability of azimuthal resolution due to orbital perturbations in the LBSAR.

Compared with current low-Earth-orbit synthetic aperture radar (SAR), geosynchronous SAR (GEO SAR) is featured with its ultrahigh orbit and ultralong integration time. In this letter, we answer the question whether the orbital perturbation items have effects on GEO SAR imaging during the long integration time and, if so, how they produce errors. To achieve these goals, we first develop a perturbing orbital elements errors model based on perturbation analysis. Then, we propose an accurate analytical expression of first to fourth Doppler parameters for GEO SAR and analyze the effects of perturbing orbital elements errors on the Doppler parameters. Furthermore, the relationship between the perturbing Doppler errors and the phase coherence of GEO SAR signal is deduced. Experiment and simulation results demonstrate the image defocusing caused by orbital perturbation. The conclusions are that the orbital perturbation does affect GEO SAR imaging by producing nonnegligible phase errors and that the required accuracy of the orbit determination should be at centimeter level in the radial direction.

Simulation of robotic space operations with minimum base reaction manipulator

T HE orbits of remote sensing systems usually dictate the ground resolution, area coverage, and frequency of coverage parameters. Lower altitude orbits enable a spacecraft to provide higher resolution measurements, but orbital perturbations are nonnegligible due to atmospheric drag [1]. Electric propulsion systems can be used to compensate for atmospheric drag [2]. However, low altitude orbits lack wide coverage. Strategies to overcome the lack of coverage, by maneuvering the spacecraft using electric propulsion, were proposed in the literature [3]. In this case, the thrusters are not only used to compensate the perturbations, but are also used to continuously maneuver the spacecraft to achieve given coverage requirements over a given set of sites in a given time frame. One example is to visit a set of 20 sites within a time frame tf of 50 days. However, this solution complicates the satellite system. The motivation then is to find a lower cost solution for this kind of mission. This note proposes a new solution in which the spacecraft is placed in a natural orbit such that it visits all the sites within the time frame without maneuvering. An advantage is the short time for visiting all sites. Performing continuous thrust transfers to maneuver the satellite from one site to another often takes more time. The disadvantage is that some sites may not be visited accurately. A tradeoff between the accuracy and the revisit time for all the sites is needed when planning a mission. To find this natural orbit, we use a genetic algorithm (GA) to perform a directed search among all possible orbits. Implementing a genetic algorithm does not guarantee an optimal orbit, however, it has been shown that in subsequent iterations, better solutions will be sampled at exponentially increasing rates [4]. This issue will be discussed in the section titled Genetic Algorithms. GAs have been adopted in the literature to solve several orbital mechanics problems. We chose to solve the ground surveillance problem using genetic algorithms because this problem is characterized by many local minima. Conventional optimization methods (e.g., gradient methods) are not suitable for this kind of problem. GAs use random choice as a tool to guide a highly exploitative search in the design space [5]. Enumerated methods scan the whole domain andfind the optimal solution. Enumeratedmethods can provide good solutions to the problem.However, the efficiency of these algorithms is very low compared with genetic algorithms [5]. TheGA solution is used as a starting point to find a local minimum solution through traditional optimization methodologies. The final solution will be a local minimum in the neighborhood of the GA solution. Two types of constraints are considered. Thefirstmission searches for maximum resolution for each site for a given imaging sensor. The second mission tries to maximize the observation time.

Microscope is a satellite due to be launched in March 2008 (Microscope is undertaken jointly by the Centre National d'Etudes Spatiales, France, the ESA, the Office National d'Etude et de Recherche en Aerospatial, France, and the Cote d'Azur Observatory, France). It has the mission of testing the equivalence principle, which postulates the equality between gravitational mass and inertial mass. This paper addresses the fault diagnosis task of the Microscope thrusters. The faulty situations correspond to thrusters blocking themselves and/or closing when in operation. Two model-based diagnosis schemes based on H ∞ /H_ filters are proposed and compared with each other in terms of performances and complexity. Nonlinear simulations show that both schemes are able to detect and isolate faults, despite the presence of measurement noises, measurement delays, sensor misalignment phenomena, and disturbances (i.e., third-body disturbances, J 2 disturbances, atmospheric drag, and solar radiation pressure).

Tethered CubeSat systems are characterized by momentum exchange and gravity gradient stabilization, and encounter stability challenges, especially with shorter tethers in highly inclined and eccentric orbits. This study investigates the libration dynamics of a 12U CubeSat in its stowed configuration, which separates into two satellites connected via tether, throughout its deployment, station-keeping, and retrieval phases. Two deployed configurations, symmetrical 6U-6U, and asymmetrical 8U-4U connected by a non-conductive tether with a maximum length of 100m are analyzed. The study accounts for perturbations including Earth’s oblateness, atmospheric drag, solar radiation pressure, and lunisolar gravitation, modeling the tether as a rigid, extendable rod and the satellites as lumped masses. The translational and rotational dynamics are decoupled, assuming the system’s center of mass follows an unperturbed Keplerian Sun-synchronous orbit at altitudes between 400 and 600km. A tension control strategy based on Lyapunov’s direct method and supplemented by external actuation torques is explored. Results show high orbit eccentricity significantly affects the maximum in-plane libration angle. The 6U-6U system experiences smaller disturbance torques but greater tether tension variations than the 8U-4U system. Both configurations exhibit larger in-plane oscillations compared to near-zero out-of-plane oscillations, nonetheless, eclipse passages exacerbate the out-of-plane libration of the 8U-4U system. Relative stability is maintained during deployment, however, retrieval is chaotic, with notable oscillations in libration angles and tether tension. The tension control strategy effectively dampens oscillations during retrieval but loses effectiveness as tether tension approaches zero at retrieval’s conclusion. External actuation torques, ranging from 1-2mNm for deployment to 15mNm for retrieval, complement tension control.