The primary relative navigation system of the PRISMA technology demonstration mission is based on GPS. The PRISMA satellites have been launched in June 2010 and, since their separation in August, have been successfully conducting a multitude of autonomous formation flying and on-orbit servicing experiments. The onboard navigation system relies on GPS measurements to provide in real-time accurate absolute and relative navigation. The output of the GPS-based navigation system is used to feed the numerous on-board feedback controllers and the safety monitoring and collision avoidance functions. As a consequence the navigation system has to provide the utmost accuracy to the experimental payloads and, at the same time, be robust and reliable to support fault detection isolation and recovery tasks. The commissioning of the GPS
navigation system is therefore a crucial milestone for the mission. It has to be ensured that the onboard navigation solution is always reliable even in the presence of hardware anomalies and that the navigation performance satisfies the mission requirements. The performance assessment is done using precise orbit products generated by a ground facility for GPS-based precise orbit determination which needs, in turn, to be also commissioned. The paper summarizes the commissioning activities performed to validate both precise orbit products and onboard navigation and presents key flight results to illustrate the overall system performance.
This paper describes the TerraSAR-X / TanDEM-X formation control concept, which bases on the relative eccentricity / inclination vector separation method, implemented within GSOC’s flight dynamics system. The paper specifically elaborates on the results obtained during the first six month of the formation flying mission showing the safe and precise automated ground-control of the space-borne radar interferometer.
The characteristics of the resonant disturbing function for an asteroid perturbed by a planet in circular orbit are discussed. The location of the libration centers and their dependence with the orbital elements of the resonant orbit are analyzed. A proposed numerical method (Gallardo 2006a) for computing the strengths of the resonances is revised and applied to the region of the main belt of asteroids showing the relevance of several mean motion resonances (MMR) with several planets.
The solar electric propulsion could be the best option for the transports of the future due to its high specific impulse when compared to the chemical propulsion. Electric propellants are being extensively used to assist the propulsion of terrestrial satellites for the maneuvers of orbit correction and as primary propulsion in missions toward other bodies of the solar system. In this work the optimization of interplanetary missions using solar electric propulsion (SEP) and Gravity Assisted Maneuver to reduce the costs of the mission, is considered. The high specific impulse of electric propulsion makes a Gravity Assisted Maneuver 1 year after departure convenient. Missions for several Near Earth Asteroids will be considered. The analysis suggests criteria for the definition of initial solutions demanded for the process of optimization of trajectories. Trajectories to the asteroid 2002TC70 are analyzed. Direct trajectories, trajectories with 1 gravity assisted at the Earth and with 2 gravity assisted with the Earth and either Mars are presented. Shall be analyzed missions with thrusters PPS1350 and the Phall 1 for performance comparison. An indirect optimization method will be used in the simulations.
In this work the problem of two-impulsive orbital transfers between non-coplanar circular or elliptical orbits is studied using hyperbolic orbits as the transfer orbits, with minimum fuel consumption but with time limit for this transfer. The equations presented by Eckel and Vinh (1984) was used. These equations provide the elliptical transfer orbit between non-coplanar elliptical orbits with minimum fuel and fixed time of transfer; or minimum time of transfer for a prescribed fuel consumption, using elliptic transfer orbits. But, in this work, only the problem with minimum fuel consumption and fixed time of transfer was considered. Then, the equations presented by Eckel and Vinh (1984) was adapted to consider the problem of non-coplanar orbital transfer between circular and elliptical orbits using hyperbolic orbits as the transfer orbit and we developed a software for orbital maneuvers. The original method, developed by Eckel and Vinh, was presented without numerical results in that paper. Thus, the modifications considering the maneuvers between circular orbits, the implementation for the hyperbolic case and the solutions using this method are contributions of this work. The software was tested simulating real maneuvers with success.
space mission concepts based on satellite formations feature several interesting and demanding design requirements. During formation deployment, re-sizing and re-orientation, the spacecraft must reach their desired positions without incurring in collisions or interfering with each other. Conventional GNC algorithms cannot support the degree of flexibility, the accuracy in the relative positioning and the fine pointing attitude requirements of future missions. The potential function method, based upon Lyapunov’s second theorem on stability, brings the advantage of robustness and flexibility, along with a light workload for the control system. The aim of this paper is to foster the understanding of existing algorithms based on this method. In particular, the attention is focused on reducing the fuel consumption through the parametric optimization of the Lyapunov function that defines the control algorithm.
Evolution of MEMS technology and increasing application of inertial and magnetic sensors in Consumer Electronics market has lead to the development of very low cost sensors. However, these sensors also present lower performance, mainly related to calibration errors and noise. The purpose of this work is to present a self-calibration algorithm for accelerometers and magnetic sensors used in a low cost attitude determination system. The proposed method evaluates errors of scale factors, biases and misalignment angles and combines the minimum variance estimator modeled by Lötters with the sensor model proposed by Foster. The algorithm was initially tested through simulation leading to good results comparing real and estimated parameters and then evaluated with the QUEST algorithm, allowing measurement of the impact of calibration errors in final calculated attitude. After calibration, maximum errors were reduced by more than 90%.
Two extensions of the fast and accurate special perturbation method recently developed by Peláez et al. are presented for elliptic motion. A comparison with Peláez's method and with the very efficient Stiefel-Scheifele's method, for the problems of oblate Earth plus Moon and continuous radial thrust, shows that the new formulations can appreciably improve the accuracy of Peláez's method and have a better performance of Stiefel-Scheifele's method. Future work will be to include the two new formulations and the original one due to Peláez into an adaptive scheme for highly accurate orbit propagation.
A star-sensor-based attitude determination system estimates spacecraft attitude with respect to the inertial coordinate system. When a satellite equipped with the star-sensor-based attitude determination system implements geocentric pointing, it requires the satellite's position information relative to the Earth's center. Conventional Earth-pointing satellites carried only low-order on-board orbit models because of their Earth-sensor-based attitude determination systems and moderate attitude accuracy requirements. The Advanced Land Observing Satellite (ALOS), which has stringent attitude control accuracy and attitude stability requirements and therefore a star-sensor-based attitude determination system, carries a high-order precision on-board orbit model that has 15 parameters and accomplishes the required accuracy by frequent updates of those parameters. This paper presents the design of the on-board orbit model and its flight results. In addition, the attitude control errors and attitude stability performance that were induced by the attitude control reference generated by the orbit model are assessed.
This work employs multiple GNSS antennas for providing short baseline readings, which are used as measurements for implementing a Kalman Filter based AHRS (Attitude and Heading Reference System). Besides eliminating the necessity of using magnetometer, which is well known to be a sensor difficult to calibrate and prone to interference, for conveying yaw information, the GNSS measurements also aid in estimation of the roll and pitch angles. For achieving high accuracy in baseline determination, GNSS carrier phase data must be used and the associated integer ambiguity problem must be solved on the fly. The proposed AHRS also incorporates gyros and accelerometers, with the gyros being employed for enabling the vehicle attitude propagation. The use of accelerometers aims at maintaining good estimation of the vehicle roll and pitch angles, in case the GNSS signal is temporarily lost. Simulated and real data are used for performance evaluation.
The purpose of this paper is to analyze the effect of albedo interference on solar sensor measurements in the Survival control mode of Brazilian remote sensing satellite Amazonia-1. Three basic algorithms for coarse Sun vector determination are investigated: the first one selects a subset of sun sensors and ignores albedo effects; the second algorithm also ignores albedo effects but uses the whole set of solar sensors, while the third one takes into account the albedo effects. A possible advantage of increasing the number of solar sensors is investigated by considering an additional set of six solar sensors in cubic configuration. Albedo model follows the 2005 annual average and standard deviation taken from Total Ozone Mapping Spectrometer TOMS Project diffuse reflectivity database, which divides the Earth surface in a set of cells with different reflectance levels. The analysis is mainly based on simulation and includes sensor noise effects. Test scenarios consider a spacecraft in LEO, Sun synchronous orbit at different altitudes, with different local time passes, different albedo reflectance levels, and different spacecraft attitudes. For every test scenario the error in the Sun vector determination is evaluated through one orbit period for different seasons and different GMT. Real data from previous satellites are considered as a model validation source.
The constantly increasing growth of the space debris population is also causing that more and more devices are looking into the sky in search of undetected objects. The process of orbit determination and further object cataloguing requires the initialisation of the object orbital state. This process is particularly complex in the cases when only angular observations from passive devices are available (e.g. topocentric right ascension and declination from a ground telescope). This paper describes the process of initial orbit determination when only a limited number of angular observations are available. Different orbital scenarios (e.g. LEO, MEO, GEO) are analysed together with the available algorithms. The analysis focuses on the suitability of the algorithm for the different orbital regimes and also in the robustness of the solution. The main objective of the analysis is to evaluate the adaptation of the algorithms and their parameterisation for the implementation in operational automated scenarios.
A spacecraft attitude estimation approach based on the Unscented Kalman Filter is derived. For nonlinear systems the Unscented Kalman Filter uses a carefully selected set of sample points to map more accurately the probability distribution than the linearization of the standard Extended Kalman Filter, leading to faster convergence from inaccurate initial conditions in attitude estimation problems. The filter formulation is based on standard attitude-vector measurements using a gyro-based model for attitude propagation. This paper compares the performance of a new technique, the Unscented Kalman Filter, when two different mathematical constructs are used to represent the attitude: the Euler angles and quaternions. In this study, the attitude of satellite is estimated with real time algorithms using real data supplied by gyros, Earth sensors and Sun sensors that are on board of the CBERS-2 (China Brazil Earth Resources Satellite).
With an aim on mission design for artificial satellites we revisit the long-term dynamics of an orbiter about a planetary moon. The main dynamical effects are modeled with the Hill-oblate problem, which is doubly averaged over the mean anomaly and the argument of the node in the rotating frame. The averaging is performed with Deprit's perturbation method and is carried up to the order three in a small parameter proportional to the ratio orbiter's mean motion system's rotation rate. However, we do not claim to have a complete third order theory because of the unconventional sorting of the Hamiltonian terms that we choose in order to extend the validity of the theory for high-altitude orbiters. The doubly averaged problem shows the existence of circular and eccentric frozen orbits, whose initial conditions in the original (non averaged) model are recovered through explicit transformation equations. In addition, control strategies based on stable-unstable manifold tours are easily designed in the averaged problem. Application to the Jupiter-Europa system illustrates the full procedure.
In this work, the resonance problem in the artificial satellites motion is studied. The development of the geopotential includes the zonal harmonics J 20 and J 40 and the tesseral harmonics J 22 and J 42 . Through succes-sive Mathieu transformations, the order of dynamical system is reduced and the final system is solved by numerical integration. In the simplified dynamical model, two critical angles are studied, φ 2201 and φ 4211 . Numerical results show the time behavior of the semi-major axis and φ 2 angle.
Effects due to resonances in the orbital motion of artificial satellites disturbed by the terrestrial tide are analyzed. The nodal co-rotation resonance, apsidal co-rotation resonance and the Lidov-Kozai's mechanism are studied. The effects of the resonances are analyzed through the variations of the metric orbital elements. Libration and circulation motions for high orbits with high eccentricities are verified for the Lidov-Kozai's mechanism.
The subject of this paper is the investigation of the effects of aerodynamic forces on the formation flight of satellites and the assessment of using differential drag as a means for formation maintenance. In order to study spacecraft formation flight in low Earth orbits, it is essential to consider the effect of J 2 . Schweighart and Sedwick have presented a set of linearized differential equations of motion, an extension of the Clohessy-Wiltshire equations, which includes the J 2 effects and are valid for circular orbits. In this paper, the Schweighart-Sedwick equations are modified to include the effects of atmospheric drag while maintaining their linearity and simplicity. These ideas are then extended to include reference orbits of small eccentricity. This results in a set of linear differential equations of motion which takes into account J 2 and drag perturbations as well as the elliptic nature of the reference orbit. Numerical simulation results based on the model are presented which characterize the effects of drag on projected circular type formations in the context of two different formation flying missions. It is shown that it is feasible to use differential drag to maintain a projected circular formation through simple panel rotation schemes.
The paper presents a combined approach to teaching of Spaceflight Dynamics, Theoretical Mechanics and Control Theory. In addition to a standard lecture course, this approach provides hands-on training at the laboratory, offering to the student a possibility to develop, design, assemble and test his/her own project of a nanosatellite attitude control system. We describe the software for mathematical modeling, the facilities for laboratory simulation of controlled system dynamics and student training based on these facilities.
The original attitude monitoring software of the EUMETSAT Polar System Flight Dynamics Facility allowed the computation of mean attitude misalignment and AOCS sensor biases in a given processing time, assuming a nominal attitude law with constant misalignments. This system has been enhanced, computing the actual attitude evolution from the angular velocity reconstructed from the gyro data and estimating all the attitude and sensor parameters with a batch least square method based on the observations provided by the on-board optical sensors. The resultant attitude evolution can be propagated after a spectral analysis of the misalignment evolution through a Fast Fourier Transform. This software has been integrated in the Metop Flight Dynamics System and it is currently being used in routine operations to estimate the satellite short term attitude evolution (only one orbit of gyro data are daily available) providing a very useful information on the pointing stability of the platform.
The problem of orbital separation of two ionosphere sounding satellites with a passive magnetic attitude control system is studied. The objective function, i.e., the expression for the altitude of a connecting chord, is derived. Optimization of the separation point in orbit and V required for separation is carried out for two separation modes: both in case of braking that completes the orbital configuration and without braking. The numerical estimation of the duration of along-track separation process and required V is done.
This paper presents a new way to deal with attitude modeling. It is designed to be able to handle highly accurate models with very complex settings while remaining simple to use, even for non attitude specialists. The representation is split in several layers, for greater flexibility and thereby provides an extensive range of attitude profiles. The core layer is a basic attitude mode similar to what can be found in many attitude simulation tools. Additional layers can be stacked to it, so that additional constraints can be taken into account, and modify the basic attitude (offsets, spin, …) It is also possible to set up an attitude sequence containing several modes linked together, in which switching from one mode to the next is events-based. Events are triggered automatically and their occurring time is computed on the fly during simulation: the user does not need to know in advance when they are supposed to occur. Numerous predefined events exist, like eclipse entry/exit, field of view entry/exit, orbital events, etc.
Systems for satellite attitude control are usually designed based on modern control techniques, which assume that the plant, sensors and actuators have linear behavior. A reaction wheel is an actuator that produces a torque as a function of electric current applied to a brushless DC motor. Ideally the output torque should be proportional to the current, but Coulomb and static frictions in the bearing introduce non linearities in the output torque, as function of the angular velocity of the wheel, especially at low speeds, near zero. These non linearities become more relevant specially in reaction wheels that are designed to rotate in both directions, causing the controller error to increase significantly at sense of speed reversals (zero crossings). This article presents results of a control orientation of an air bearing table by means of a reaction wheel. To check the controller action during a zero crossing, a small fan was attached to the table, producing a torque whose magnitude can be altered by adjusting the direction of the airflow. The control loop uses a fiber optic gyroscope (FOG) as angular velocity sensor. A PID digital controller drives the wheel based on the angular position of the table with respect to a given reference. In order to have a controller with a linear output torque, a mathematical model of torque was developed as function of the input current and wheel speed, from which it was constructed an algorithm of the inverse function, so that the non linearities were partially compensated (dynamic compensation). The results indicated that dynamic compensation can effectively reduce the maximum pointing error during zero crossing, while keeping constant other parameters such as response and settlement time.
Attitude Control System (ACS) for flexible space satellites demands great reliability, autonomy and robustness. These flexible structures face low stiffness due to minimal mass weight requirements. Satellite ACS design usually based on computer simulations without experimental verification can face instability and/or inefficient controller performance due to model uncertainties. In this paper one investigates the robustness and performance of the time domain approach LQG (Lineal Quadratic Gaussian) and the frequency domain H– Infinity approach. The satellite ACS design is performed initially in a computer simulation environment, following experimentally verification of the same control algorithm, using Quanser rotary flexible link module. This investigation has shown that the controller performance based on simulation model can be degraded when applied in an experimental set up. So this prototype verification is fundamental before satellite onboard computer algorithms implementation.
Numerical explorations show how the known periodic solutions of the Hill problem are modified in the case of the attitude-orbit coupling that may occur for large satellite structures. We focus on the case in which the elongation is the dominant satellite's characteristic and find that a rotating structure may remain with its largest dimension in a plane parallel to the plane of the primaries. In this case, the effect produced by the non-negligible physical dimension is dynamically equivalent to the perturbation produced by an oblate central body on a masspoint satellite. Based on this, it is demonstrated that the attitude-orbital coupling of a long enough body may change the dynamical characteristics of a periodic orbit about the collinear Lagrangian points.
This paper presents an overview of the attitude ground system (AGS) currently under development for the Magnetospheric Multiscale (MMS) mission. The primary responsibilities for the MMS AGS are definitive attitude determination, validation of the onboard attitude filter, and computation of certain parameters needed to improve maneuver performance. For these purposes, the ground support utilities include attitude and rate estimation for validation of the onboard estimates, sensor calibration, inertia tensor calibration, accelerometer bias estimation, center of mass estimation, and production of a definitive attitude history for use by the science teams. Much of the AGS functionality already exists in utilities used at NASA's Goddard Space Flight Center with support heritage from many missions, but new utilities are being created specifically for the MMS mission, such as for the inertia tensor, accelerometer bias, and center of mass estimation. Algorithms and test results for all the major AGS subsystems are presented here.
Passive attitude control systems are discussed. Attention is paid to the gravitational attitude control systems and, specifically, to the problem of efficient damping of the small oscillations in the vicinity of the equi-librium position in a circular orbit. The problem of determining the maximal damping rate is considered, and corresponding numerical and analytical methods are discussed. The results of optimization of several specific systems are reviewed. The analytical approach is extended to investigate the satellite-stabilizer system in a spe-cial case when there is no spring resistance in the hinge connecting the bodies.
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.
This paper presents a performance analysis of an autonomous orbit control procedure using a simplified GPS navigator (Galski et al., 2001), where the ground track drift of the satellite is estimated on-board with help of a recently developed approach (Orlando et al., 2009) that directly calculates the acceleration of the orbit ground track as a function of the solar and geomagnetic activity. The simplified navigation procedure improves the coarse geometric navigation solution provided by GPS receivers. This is done by using the GPS solutions as inputs (observations) for a real time Kalman filtering process. The orbital state vector is extended and includes the systematic error imposed to the GPS geometric solution by the changes in the set of satellites which are visible to the receiver. The simplified navigator has reduced computational cost, allowing it to be carried and executed on-board of spacecrafts. The improved outputs of this process are used in the computational implementation of an autonomous control system for the ground track drift of the spacecraft orbit. The behavior of the system is evaluated by means of orbit simulations using a CBERS-like phased remote sensing satellite. The aim of the paper is to verify if the coupled system is able to correctly calculate and perform variable size semi-major axis orbit increment maneuvers in order to keep the satellite ground track within its allowed limits (±4km).
The Meteosat Third Generation (MTG) mission requires two 3-axis stabilized satellites, one implementing the imaging instrument (MTG-I) and another the sounding instrument (MTG-S), to be operated in orbit simultaneously in collocation around the 0º geostationary position. Satellites collocation, even if requiring more complex operations, may allow operating both satellites using a single S-band station, with an important reduction of the ground support required for Flight Dynamics operations, and consequently of the number of needed ground stations and of the operational workload. Further reduction of the operational workload, together with a potential benefit in terms of availability of the scientific data, could be achieved if the collocation strategy is implemented through very large inclination maneuvers executed yearly.The paper will present in detail the analyses performed to assess the flight dynamics feasibility of the proposed operational strategy and will address its potential benefit; at the same time the major constraints at system level, which currently may prevent the implementation of the described strategy, are presented.
Diamond-like Carbon (DLC) films have been the focus of extensive research in recent years due to their potential application as surface coatings. In this paper, we report the main results obtained in our laboratories from the production and characterization of nanocrystalline diamond particles (NCD) incorporated in DLC films for biomedical applications. The films were growth on 316L stainless steel substrates from a dispersion of NCD nanoparticles in hexane using plasma enhanced chemical vapor deposition. Raman scattering spectroscopy was used to study the atomic arrangements of the film, and atomic force microscopy, the roughness of the film. The investigation of NCD-DLC electrochemical corrosion behavior was performed using potentiodynamic method. Cell viability was evaluated with L-929 mouse fibroblast cells using 2-(4,5-dimethyl-2-thiazolyl)-3,5-diphenyl-2H-tetrazolium bromide (MTT) in vitro assay. The NCD particles increased the structural diamond-like domains and surface roughness, which improved DLC and stainless steel electrochemical corrosion resistance and prevented aggressive ions from attacking metallic surfaces. Those NCD particles also increased the DLC cell viability, maximizing the potential use of NCD-DLC films in biomedical applications.
This paper presents and discuss some possibilities concerning to the design of a Regional Positioning System -RPS satellite constellation to cover the Brazilian territory. The paper presents an overview of constellation design possibilities for a hypothetical Brazilian Regional Positioning System. Since Brazil is located near the equator plane, the idea is to think about a RPS satellite constellation that takes advantage of this fact in order to design a cost effective regional system that aims at covering primarily only the Brazilian territory. In this sense, three satellite constellation types were considered. The first one is composed by Low Earth Orbit -LEO satellites with low to moderate orbital plane inclination angles. The second one is based on Medium Earth Orbit -MEO satellites placed in the equator plane or near to it, and the third one is based only on geosynchronous satellites. A preliminary optimization process was run in order to obtain an initial guess of a good design. The obtained design constellations are presented and discussed.
This paper describes the different changes implemented in a conjunction assessment and collision risk evaluation tool with the aim of reducing drastically the computational cost to ensure that a scenario where all space debris objects are analyzed against each other can be carried out in a short period of time. Improvements at algorithm level and parallelization techniques are used to shorten the time needed for the process of conjunction assessment. In the case of the collision risk evaluation, an approach for the propagation of the state covariance is presented based on the Simplified General Perturbations theory commonly used to propagate Two Line Elements.
This work shows the results of a procedure of trajectory computation through integrating DGPS/SBAS and IMU. The GPS navigator together with the differential corrections transmitted by geostationary satellite are integrated to the aircraft IMU -Inertial Measurements Unit. The system design, software development and integration, experiments and results are described for critical conditions.
This paper presents a new way to deal with transition matrices handling in variational equations. While working simultaneously on several problems dealing with orbit propagation: low-thrust trajectories and orbit determination, we implemented a feature allowing to add user equations to a propagator in order to solve the first problem. Then it appeared that this feature could be reused to deal with the second one. Indeed, the orbit determination problem is based on variational equations involving transition matrices, which can also be considered as additional parameters, propagated at the same time as the original state vector. This method allows a very modular implementation of both problems, with looser coupling in the equations. It has been successfully integrated in the Orekit open-source library.
As a multi-mission GEO satellite, COMS has three payloads including Ka-band communications, geostationary ocean color imager, and meteorological imager. COMS flight dynamics system provides the general on-station functions such as orbit determination, orbit prediction, event prediction, station-keeping maneuver planning, station-relocation maneuver planning, and fuel accounting. There are some specific flight dynamics functions to operate the Eurostar 3000 spacecraft bus such as wheel off-loading management, oscillator updating management, and on-station attitude reacquisition management. In this paper, operational validation of the major functions in COMS flight dynamics system is presented for the first six month of operations period.
One of the measures of guidelines for space debris (SD) mitigation in the maintained near-earth space orbits is a removal of the man-made SD, in particular these ones which have terminated their mission plans. For the low-earth orbits it means the transfer of SD objects into the new (so-called disposal) orbits having lifetime not more than 25 years. The task of a search for the so-called disposal orbits, which can ensure the given limited lifetime for the started to move in them space objects (SO), and to which it is necessary to re-orbit the SD objects being in near-polar orbits with the mean altitudes Н ср ~ 750-850 km is considered. The disposal orbit is formed on the basis of an original orbit by the application in some its point of a corresponding impulse of velocity ΔV, directed against the space vehicle' motion. At an estimation of a required fuel content for a realization of the re-orbital maneuver it is supposed, that this maneuver implements by means of a propulsion system having a specific impulse J sp =360s, and the initial mass of a space craft m 0 ~ 500kg. As a result the estimates of lifetime of SO, starting orbiting in the disposal orbits and recommendations for selection of the disposal orbits for the examined class of the near-polar orbits ensuring the limited lifetime for de-orbited space objects are made.
We present the guidance approach for the Small Solar Power Sail Demonstrator, IKAROS. This approach will be used for future solar sail missions. First, we show the range of orbital controllability, which is severely restricted by the attitude constraint from the power and communication. Second, we also present the adaptive guidance method, in which guidance parameters are updated based on the estimation of the spacecraft's parameters such as the area and reflectivity coefficient of the sail. Finally, we evaluate the guidance approach performance by comparing to flight data of IKAROS.
In the present work, the influence of the Sun on the fuel consumption of transfers from circular low Earth orbits (LEOs) to circular low Moon orbits (LMOs) is investigated. The class of two impulse trajectories is considered: a first accelerating velocity impulse tangential to the space vehicle velocity relative to Earth is applied at a circular low Earth orbit and a second braking velocity impulse tangential to the space vehicle velocity relative to Moon is applied at a circular low Moon orbit. The fuel consumption is equivalent to the total characteristic velocity which is defined by the arithmetic sum of velocity changes. Local optimal transfers are calculated through two different approaches: inner transfers and Belbruno-Miller transfers. In both cases, the optimization problem is solved by means of an algorithm based on gradient method in conjunction with Newton-Raphson method.
We consider the motion of a dumbbell in a central Newtonian field of gravity. The dumbbell consists of two massive points connected by a massless rod. The third massive point, a cabin, slides along the rod according to a prescribed rule. One can consider such a mechanical system as a simplified model for an orbital tethered elevator. We study the most interesting case when the elevator cabin moves periodically between the dumbbell end masses. For the sake of simplicity, only planar motions of the system are analyzed. Assuming the cabin mass small compared to the masses of the endpoints, we use the Poincaré theory to determine the conditions of exis-tence for families of periodic motions that depend analytically on the appropriate small parameter. These mo-tions tend to corresponding stable radial relative equilibria as the small parameter tends to zero. Each one of the two existing relative equilibria generates exactly one family of such periodic motions if the parameter is small enough. Stability of these periodic motions is investigated in the linear approximation. The solutions are computed up to the terms of the first order with respect to the small parameter. Keywords: orbital dumbbell, tethered satellite system, space elevator, Poincaré's theorem on periodic motions, stability of periodic motions. 1 Problem formulation and main notations We study dynamics of the massive points M 0 , "space station", M 1 , "sub-satellite", and M 2 , "elevator's cabin", in field of gravity generated by an attracting center P, "planet", fixed in the absolute space (see Fig. 1). Denote m 0 , m 1 , and m 2 masses of the points M 0 , M 1 , and M 2 respectively. Points M 0 and M 1 are connected by a massless rod, and the point M 2 slides along this rod according to a given rule, e.g., it can move periodically between M 0 and M 1 . Within this simplified approach this system describes dynamics of the tethered dumbbell satellite with an elevator if the tension of the tether is sufficiently large and the mass of the cabin is sufficiently small. The above holds if the system composed by the station and the sub-satellite is close to the stable radial equilibria, found and studied earlier in Beletsky and Ponomareva (1990), Kosenko and Stepanov (2006). For such configurations the tether is stretched, so it can be correctly approximated by a rod. Then the whole system can be considered as a large-scaled orbiting object equipped with a space elevator.
The study presented in this paper deals with Geostationary Transfer Orbits (GTO) for which new French regulations (defined in the context of the French Space Act) will fully apply at the end of 2011. Geostationary Transfer Orbits are characterized by a low perigee (altitude of a few hundreds of kilometres) and a high apogee (altitude typically identical to that of geostationary satellites) among other features. The objective of the study is to analyze the dynamics of objects in geostationary transfer orbits in order to better understand what the lifetime (time during which the object remains in orbit) most depends on. Because of the high eccentricity, the orbit is strongly affected by the gravitational effects of the Sun and Moon. But because the perigee is low, drag has a strong impact too. The coupling of the two perturbations combined with the effects of the Earth potential (secular drifts mainly) makes the orbit's evolution particularly sensitive to initial conditions and modelling errors. One key element is the initial position of the Sun (and to a lesser extent the Moon) which changes the mean altitude of the perigee, which translates into more or less drag, hence more or less decrease rate of the semi-major axis at the beginning of the lifetime. But when the semi-major axis reaches a value of around 15000km, the perigee altitude may increase or decrease strongly because the angle between the Sun and the line of apsides is then nearly constant. The paper attempts to explain all these aspects and discusses the possibility of limiting the lifetime of objects in Geostationary Transfer Orbits.
This article investigates the use of a multilayer feedforward artificial neural network into a GPS integrated low cost inertial navigation system based on MEMS sensors. The neural network is applied as an alternative of integration technique, with the purpose of providing better navigation solutions, during the lack of information in GPS outages portions of time. An input-output neural network signals model is proposed, based on a set of simplified terrestrial vehicle navigation equations. Also an adaptive Kalman filter training methodology is tested with real navigation data. Preliminary simulated numerical results are presented, based on urban vehicular positioning application data trials, acquired from low cost Crossbow CD400-200 IMU and an Ashtech Z12 GPS receiver.