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Thesis (M.Com.(quantitative Management)) / Satellite Scheduling Problems (SSP) are NP-hard and constraint programming and metaheuristics solution methods yield mixed results. This study investigates a new version of the SSP, the Satellite Constellation Time-Window Optimization Problem (SCoTWOP), involving commercial satellite constellations that provide frequent earth coverage. The SCoTWOP is related to the dual of the Vehicle Routing Problem with Multiple Timewindows, suggesting binary solution vectors representing an activation of time-windows. This representation fitted well with the MatLab® Genetic Algorithm and Direct Search Toolbox subsequently used to experiment with genetic algorithms, tabu search, and simulated annealing as SCoTWOP solution methods. The genetic algorithm was most successful and in some instances activated all 250 imaging time-windows, a number that is typical for a constellation of six satellites. / Quantitative Management

Στην παρούσα διατριβή ασχολούμαστε με την μελέτη περιοδικών και ασυμπτωτικών λύσεων στο περιορισμένο πρόβλημα των τεσσάρων σωμάτων. Πιο συγκεκριμένα: Στο κεφάλαιο 1 περιγράφουμε το πρόβλημα των τριών και των τεσσάρων σωμάτων, κάνοντας μια ιστορική αναδρομή και παραθέτουμε τις αρχικές εξισώσεις της κίνησης. Στο κεφάλαιο 2 μελετάμε αριθμητικά το περιορισμένο πρόβλημα των τεσσάρων σωμάτων, στην Lagrangian διαμόρφωση. Υπολογίζουμε τα σημεία ισορροπίας, καθώς και τις επιτρεπτές περιοχές κίνησης του τέταρτου σώματος. Στο κεφάλαιο 3 μελετάμε την ευστάθεια των σημείων ισορροπίας. Επίσης υπολογίζουμε και παρουσιάζουμε τις περιοχές έλξης, για το δυναμικό σύστημα των τεσσάρων σωμάτων. Στο κεφάλαιο 4 μελετάμε οικογένειες απλών συμμετρικών και μη συμμετρικών περιοδικών τροχιών του περιορισμένου προβλήματος των τεσσάρων σωμάτων. Υπολογίζουμε για κάθε περίπτωση τιμών των μαζών, σειρές κρίσιμων περιοδικών τροχιών κάθε οικογένειας ξεχωριστά. Τέλος στο κεφάλαιο 5 μελετάμε αριθμητικά οικογένειες απλών ασύμμετρων περιοδικών τροχιών στο περιορισμένο πρόβλημα των τεσσάρων σωμάτων, έχοντας θέσει ως πρωτεύοντα σώματα τους ΄Ηλιο - Δία και έναν Τρωικό Αστεροειδή και θεωρώντας ως τέταρτο αμελητέας μάζας σώμα ένα διαστημόπλοιο. Τα πρωτεύοντα σώματα υπακούουν στην ευσταθή Lagrangian τριγωνική διαμόρφωση. Μελετήσαμε επίσης αναλυτικά και αριθμητικά τις λύσεις στην περιοχή των ευσταθών σημείων ισορροπίας του συστήματος, βρήκαmε οικογένειες περιοδικών λύσεων και μελετήσαμε την γραμμική ευστάθεια τους. Τα αποτελέσματα των κεφαλαίων 2,3,4 και 5 έχουν δημοσιευτεί σε τρία διεθνή περιοδικά και ένα κομμάτι του κεφαλαίου 5 παρουσιάστηκε σε διεθνές συνέδριο (με συγγραφείς τους Μπαλταγιάννη Α. και Παπαδάκη Κ.). Πιο συγκεκριμένα η μελέτη των κεφαλαίων 2 και 3 έχει δημοσιευτεί στο περιοδικό “International Journal of Bifurcation and Chaos, 21, 2011, pp. 2179-2193” με τον τίτλο: “Equilibrium Points and their stability in the restricted four-body problem”. Τα αποτελέσματα του κεφαλαίου 4 δημοσιεύτηκαν mε τον τίτλο: “Families of periodic orbits in the restricted four-body problem” στο περιοδικό “Astrophysics and Space Science, 336, 2011, pp. 357-367”. Επίσης το κεφάλαιο 5 υπό τον τίτλο “Periodic solutions in the Sun - Jupiter - Trojan Asteroid - Spacecraft system”, δημοσιεύτηκε στο περιοδικό ”Planetary and Space Science, 75, 2013, pp. 148-157”. Το διεθνές συνέδριο στο οποίο παρουσιάστηκε τμήμα του κεφαλαίου 5 ήταν το : “10th Hellenic Astronomical Conference, Proceedings of the conference held at Ioannina, Greece, 5-8 September 2011, pp. 23-24” και η εργασία είχε τίτλο: “Families of periodic orbits in the Sun - Jupiter - Trojan Asteroid system”. Η παρούσα διατριβή εκπονήθηκε με την οικονομική υποστήριξη του ερευνητικού προγράμματος του Πανεπιστημίου Πατρών: Κ. Καραθεοδωρή. / In this thesis we are concerned with the periodic and asymptotic solutions of the restricted four - body problem. In chapter 1 we describe the three - body and four - body problem, starting with historical information. We also present the needed equations of motion and integrals of the problem. In chapter 2 we study numerically the problem of four - bodies, according to the Lagrangian equilateral triangle configuration. We find the equilibrium points and the allowed regions of motion. In chapter 3 we study the stability of the relative equibrium solutions. We also illustrate the regions of the basins of attraction for the equilibrium points of the present dynamical model. In chapter 4 we present families of simple symmetric and non-symmetric periodic orbits in the restricted four-body problem. Series of critical periodic orbits of each family and in any case of the mass parameters are also calculated. In chapter 5 we study, numerically, families of simple non-symmetric periodic orbits of the restricted four-body problem, where we consider the three primary bodies as Sun, Jupiter and a Trojan Asteroid and as a massless fourth body, a spacecraft. The primary bodies are set in the stable Lagrangian equilateral triangle configuration. We also study analytically the solutions in the neighborhood of the stable equilibrium points and the linear stability of each periodic solution. The results of the chapters 2,3,4 and 5 have been published in three journals and a part of chapter 5 has been presented in an international conference. Chapters 2 and 3 have been published in “International Journal of Bifurcation and Chaos, 21, 2011, pp. 2179-2193” under the title of “Equilibrium Points and their stability in the restricted four-body problem”. Chapter 4 has been titled “Families of periodic orbits in the restricted four- body problem” and published in “Astrophysics and Space Science, 336, 2011, pp. 357-367”. Chapter 5 has been titled “Periodic solutions in the Sun - Jupiter - Trojan Asteroid - Spacecraft system,” and published in “Planetary and Space Science, 75, 2013, pp. 148-157”. The conference was the “10th Hellenic Astronomical Conference, Proceedings of the conference held at Ioannina, Greece, 5-8 September 2011, pp. 23-24” and part of the chapter 5 was presented under the title of “Families of periodic orbits in the Sun - Jupiter - Trojan Asteroid system”. This thesis was compiled while the author was in receipt of “K.Karatheodory” research grant.

Thesis (M.Com.(quantitative Management)) / Satellite Scheduling Problems (SSP) are NP-hard and constraint programming and metaheuristics solution methods yield mixed results. This study investigates a new version of the SSP, the Satellite Constellation Time-Window Optimization Problem (SCoTWOP), involving commercial satellite constellations that provide frequent earth coverage. The SCoTWOP is related to the dual of the Vehicle Routing Problem with Multiple Timewindows, suggesting binary solution vectors representing an activation of time-windows. This representation fitted well with the MatLab® Genetic Algorithm and Direct Search Toolbox subsequently used to experiment with genetic algorithms, tabu search, and simulated annealing as SCoTWOP solution methods. The genetic algorithm was most successful and in some instances activated all 250 imaging time-windows, a number that is typical for a constellation of six satellites. / Quantitative Management

Autonomous navigation is the determination of satellites position and velocity vectors onboard the satellite, using the measurements available onboard. The orbital information of a satellite needs to be obtained to support different house keeping operations such as routine tracking for health monitoring, payload data processing and annotation, orbit manoeuver planning, and prediction of intrusion in various sensors' field of view by celestial bodies like Sun, Moon etc. Determination of the satellites orbital parameters is done in a number of ways using a variety of measurements. These measurements may originate from ground based systems as range and range rate measurements, or from another satellite as in the case of GPS (Global Positioning System) and TDUSS (Tracking Data Relay Satellite Systems), or from the same satellite by using sensors like horizon sensor^ sun sensor, star tracker, landmark tracker etc. Depending upon the measurement errors, sampling rates, and adequacy of the estimation scheme, the navigation accuracy can be anywhere in the range of 10m - 10 kms in absolute location. A wide variety of tracking sensors have been proposed in the literature for autonomous navigation. They are broadly classified as (1) Satellite-satellite tracking, (2) Ground- satellite tracking, (3) fully autonomous tracking. Of the various navigation sensors, it may be cost effective to use existing onboard sensors which are well proven in space. Hence, in the current thesis, the Horizon scanner is employed as the primary navigation sensor-. It has been shown in the literature that by using horizon sensors and gyros, a high accuracy pointing of the order of .01 - .03 deg can be achieved in the case of low earth orbits. Motivated by such a fact, the current thesis deals with autonomous orbit determination using measurements from the horizon sensors with the assumption that the attitude is known to the above quoted accuracies. The horizon scanners are mounted on either side of the yaw axis in the pitch yaw plane at an angle of 70 deg with respect to the yaw axis. The Field Of View (FOV) moves about the scanner axis on a cone of 45 deg half cone angle. During each scan, the FOV generates two horizon points, one at the space-Earth entry and the other at the Earth-space exit. The horizon points, therefore, lie• on the edge of the Earth disc seen by the satellite. For a spherical earth, a minimum of three such horizon points are needed to estimate the angular radius and the center of the circular horizon disc. Since a total of four horizon points are available from a pair of scanners, they can be used to extract the satellite-earth distance and direction.These horizon points are corrupted by noise due to uncertainties in the Earth's radiation pattern, detector mechanism, the truncation and roundoff errors due to digitisation of the measurements. Owing to the finite spin rate of the scanning mechanism, the measurements are available at discrete time intervals. Thus a filtering algorithm with appropriate state dynamics becomes essential to handle the •noise in the measurements, to obtain the best estimate and to propagate the state between the measurements. The orbit of a low earth satellite can be represented by either a state vector (position and velocity vectors in inertial frame) or Keplerian elements. The choice depends upon the available processors, functions and the end use of the estimated orbit information. It is shown in the thesis that position and velocity vectors in inertial frame or the position vector in local reference frame, do result in a simplified, state representation. By using the f and g series method for inertial position and velocity, the state propagation is achieved in linear form. i.e. Xk+1 = AXK where X is the state (position, velocity) and A the state transition matrix derived from 'f' and 'g' series. The configuration of a 3 axis stabilised spacecraft with two horizon scanners is used to simulate the measurements. As a step towards establishing the feasibility of extracting the orbital parameters, the governing equations are formulated to compute the satellite-earth vector from the four horizon points generated by a pair of Horizon Scanners in the presence of measurement noise. Using these derived satellite-earth vectors as measurements, Kalman filter equations are developed, where both the state and measurements equations are linear. Based on simulations, it is shown that a position accuracy of about 2 kms can be achieved. Additionally, the effect of sudden disturbances like substantial slewing of the solar panels prior and after the payload operations are also analysed. It is shown that a relatively simple Low Pass Filter (LPF) in the measurements loop with a cut-off frequency of 10 Wo (Wo = orbital frequency) effectively suppresses the high frequency effects from sudden disturbances which otherwise camouflage the navigational information content of the signal. Then Kalman filter can continue to estimate the orbit with the same kind of accuracy as before without recourse to re-tuning of covariance matrices. Having established the feasibility of extracting the orbit information, the next step is to treat the measurements in its original form, namely, the non-linear form. The entry or exit timing pulses generated by the scanner when multiplied by the scan rate yield entry or exit azimuth angles in the scanner frame of reference, which in turn represents an effective measurement variable. These azimuth angles are obtained as inverse trigonometric functions of the satellite-earth vector. Thus the horizon scanner measurements are non-linear functions of the orbital state. The analytical equations for the horizon points as seen in the body frame are derived, first for a spherical earth case. To account for the oblate shape of the earth, a simple one step correction algorithm is developed to calculate the horizon points. The horizon points calculated from this simple algorithm matches well with the ones from accurate model within a bound of 5%. Since the horizon points (measurements) are non-linear functions of the state, an Extended Kalman Filter (EKF) is employed for state estimation. Through various simulation runs, it is observed that the along track state has got poor observability when the four horizon points are treated as measurements in their original form, as against the derived satellite-earth vector in the earlier strategy. This is also substantiated by means of condition number of the observability matrix. In order to examine this problem in detail, the observability of the three modes such as along-track, radial, and cross-track components (i.e. the local orbit frame of reference) are analysed. This difficulty in observability is obviated when an additional sensor is used in the roll-yaw plane. Subsequently the simulation studies are carried out with two scanners in pitch-yaw plane and one scanner in the roll-yaw plane (ie. a total of 6 horizon points at each time). Based on the simulations, it is shown that the achievable accuracy in absolute position is about 2 kms.- Since the scanner in the roll-yaw plane is susceptible to dazzling by Sun, the effect of data breaks due to sensor inhibition is also analysed. It is further established that such data breaks do not improve the accuracy of the estimates of the along-track component during the transient phase. However, filter does not diverge during this period. Following the analysis of the' filter performance, influence of Earth's oblateness on the measurement model studied. It is observed that the error in horizon points, due to spherical Earth approximation behave like a sinusoid of twice the orbital frequency alongwith a bias of about 0.21° in the case of a 900 kms sun synchronous orbit. The error in the 6 horizon points is shown to give rise to 6 sinusoids. Since the measurement model for a spherical earth is the simplest one, the feasibility of estimating these sinusoids along with the orbital state forms the next part of the thesis. Each sinusoid along with the bias is represented as a 3 state recursive equation in the following form where i refers to the ith sinusoid and T the sampling interval. The augmented or composite state variable X consists of bias, Sine and Cosine components of the sinusoids. The 6 sinusoids together with the three dimensional orbital position vector in local coordinate frame then lead to a 21 state augmented Kalman Filter. With the 21 state filter, observability problems are experienced. Hence the magnetic field strength, which is a function of radial distance as measured by an onboard magnetometer is proposed as additional measurement. Subsequently, on using 6 horizon point measurements and the radial distance measurements obtained from a magnetometer and taking advantage of relationships between sinusoids, it is shown that a ten state filter (ie. 3 local orbital states, one bias and 3 zero mean sinusoids) can effectively function as an onboard orbit filter. The filter performance is investigated for circular as well as low eccentricity orbits. The 10-state filter is shown to exhibit a lag while following the radial component in case of low eccentricity orbits. This deficiency is overcome by introducing two more states, namely the radial velocity and acceleration thus resulting in a 12-state filter. Simulation studies reveal that the 12-state filter performance is very good for low eccentricity orbits. The lag observed in 10-state filter is totally removed. Besides, the 12-state filter is able to follow the changes in orbit due to orbital manoeuvers which are part of orbit acquisition plans for any mission.

O estudo da região de estabilidade é de extrema importância nas ciências, aplicações em engenharia e nos sistemas de controle não linear. Neste trabalho, uma caracterização completa da região de estabilidade e da fronteira da região de estabilidade de pontos fixos estáveis de uma classe ampla de sistemas dinâmicos discretos não lineares é desenvolvida. Os resultados deste trabalho estendem a caracterização da região de estabilidade já proposta na literatura para uma ampla classe de sistemas, modelados por difeomorfismos e que admitem a presença de órbitas periódicas e pontos fixos na fronteira da região de estabilidade. Caracterizações dinâmicas e topológicas são propostas para a fronteira da região de estabilidade. Além disso, são dadas condições necessárias e suficientes para que um ponto fixo ou órbita periódica pertença à fronteira da região de estabilidade. Exemplos numéricos, incluindo o modelo de uma rede neural simétrica com 2-neurônios, ilustram os resultados propostos neste trabalho. / The study of the stability region is very important in the sciences, engineering applications, and in nonlinear control systems. In this work, a complete characterization for both the stability region and the stability boundary of stable xed points of a nonlinear discrete dynamical systems is developed. The results of this work extend the characterization of the stability region already proposed in the literature for a larger class of systems, which are modeled by dieomorphisms and which admit the presence of periodic orbits and xed points on the stability boundary. Several dynamical and topological characterizations are proposed to the stability boundary. Moreover, several necessary and sucient conditions for xed points and periodic orbits to lie on the stability boundary are derived. Numerical examples, including the model of a symmetric neural network with 2-neurons, illustrate the results proposed in this work.

Estudamos o comportamento dinâmico de séries temporais experimentais obtidas de um circuito de Chua quando dois parâmetros de controle, $Delta R_1$ e $Delta R_2$, são variados.Investigamos os comportamentos caótico e periódico, analisando as séries temporais ao redor e no interior de duas janelas periódicas presentes no espaço de parâmetros $(Delta R_1,Delta R_2)$ do circuito. Na vizinhança da janela de período três, analisamos como a dinâmica simbólica se altera quando construída em diferentes seções de Poincaré de um mesmo atrator, e investigamos a dimensão dos mapas de retorno, uni ou bidimensional, para diferentes atratores caóticos presentes nessa região do espaço de parâmetros. Ainda nessa vizinhança, empregamos técnicas de caracterização topológica para confirmar a existência de fibras caóticas, que são curvas de codimensão um no espaço de parâmetros onde as propriedades caóticas dos atratores são preservadas.Ao redor da janela de período quatro, investigamos a transição entre os três comportamentos caóticos para os quais construímos os respectivos moldes topológicos. Propusemos também um molde topológico para o regime caótico após a crise por fusão ocorrer no circuito. Finalizando, investigamos as bifurcações e a estrutura topológica das órbitas periódicas que formam as janelas de período três e de período quatro, construindo um espaço de parâmetros topológico, baseado em um mapa bi-modal, para descrever as duas janela periódicas. / We have studied the dynamical behavior of experimental time series obtained from a Chua's circuit by variation of two parameter control, $Delta R_1$ and $Delta R_2$. We investigated the chaotic and periodic behaviors of the circuit, analyzing temporal series around and inside of two periodic windows in the two-parameter space $(Delta R_1,Delta R_2)$. In the period-three window neighborhood, we analyzed how the symbolic dynamics changes when it is built by different Poincaré sections of an attractor, and we studied the dimension of return map, one- or two-dimensional, for many chaotic attractors in this region of the parameter space. In this neighborhood, we also applied topological techniques to confirm the existence of chaotic fibers: codimension one curves where the chaotic properties of the attractors remain unchanged in the two-parameter space.Around the period-four window, we investigated, by template analysis, the transition between three chaotic attractors found in the Chua's circuit. We proposed a template for chaotic regime of the circuit after merge-crisis. Finally, we investigated the bifurcations and topological structure of periodic orbits in period-three and period-four windows and also proposed a topological parameter space, based in a bimodal map model, that describe these two periodic windows.

O estudo da região de estabilidade é de extrema importância nas ciências, aplicações em engenharia e nos sistemas de controle não linear. Neste trabalho, uma caracterização completa da região de estabilidade e da fronteira da região de estabilidade de pontos fixos estáveis de uma classe ampla de sistemas dinâmicos discretos não lineares é desenvolvida. Os resultados deste trabalho estendem a caracterização da região de estabilidade já proposta na literatura para uma ampla classe de sistemas, modelados por difeomorfismos e que admitem a presença de órbitas periódicas e pontos fixos na fronteira da região de estabilidade. Caracterizações dinâmicas e topológicas são propostas para a fronteira da região de estabilidade. Além disso, são dadas condições necessárias e suficientes para que um ponto fixo ou órbita periódica pertença à fronteira da região de estabilidade. Exemplos numéricos, incluindo o modelo de uma rede neural simétrica com 2-neurônios, ilustram os resultados propostos neste trabalho. / The study of the stability region is very important in the sciences, engineering applications, and in nonlinear control systems. In this work, a complete characterization for both the stability region and the stability boundary of stable xed points of a nonlinear discrete dynamical systems is developed. The results of this work extend the characterization of the stability region already proposed in the literature for a larger class of systems, which are modeled by dieomorphisms and which admit the presence of periodic orbits and xed points on the stability boundary. Several dynamical and topological characterizations are proposed to the stability boundary. Moreover, several necessary and sucient conditions for xed points and periodic orbits to lie on the stability boundary are derived. Numerical examples, including the model of a symmetric neural network with 2-neurons, illustrate the results proposed in this work.

The motion of neutral test particles, both massive and massless, in the space time of a charged source described by the Reissner-Nordström solution is studied. This solution is characterized by two parameters, mass and charge, which defines the horizons of the source. When the mass is larger than the charge, the solution describes a black hole, with two distinct horizons. When the mass and charge are equal there is an extremal black hole, and both horizons merge to one. Finally, when the charge is larger than the mass there is a naked singularity, with no horizon. The structure and properties of these different type of solution are presented and discussed. A general solution of the equations of motion is presented in function of the Weierstrass elliptic function ℘. In addition, the possible orbits for test particles are discussed, and the conditions for existence of closed, circular or escape orbits are presented. The classifications is made based on the particles energy, and the mass and charge of the source. We find that all mentioned orbits are allowed for the three different type of solutions. In particular, for extremal black holes and naked singularities, we find stable circular orbits located outside the event horizon and hence being visible for an external observer. / O movimento de partículas teste neutras, ambas massivas e sem massa, no espaço-tempo de uma fonte carregada descrita pela solução de Reissner-Nordström é estudada. Essa solução é caracterizada por dois parâmetros, massa e carga, que definem os horizontes da fonte. Quando a massa é maior que a carga tal solução descreve um buraco negro com dois horizontes distintos. Quando a massa e a carga são iguais há um buraco negro extremo, e ambos os horizontes se unem em um. Finalmente, quando a carga é maior que a massa, há uma singularidade nua, sem horizontes. A estrutura e as propriedades dessas diferentes soluções são apresentadas e discutidas. Uma solução geral da equação de movimento é apresentada em termos da função elíptica de Weierstrass, ℘. Além do mais as possiveis órbitas para uma partícula teste são discutidas, e as condições para existência de órbitas fechadas, circulares e de escape são apresentadas. A classificação é feita a partir da energia da partícula, e da massa e carga da fonte. Encontramos que todas as orbitas mencionadas são permitidas nos três diferentes tipos de soluções. Em partícular, para buracos negros extremos e singularidades nuas, encontramos órbitas circulares estáveis localizadas fora do horizonte de eventos e, consequentemente, sendo visível para observadores externos.

Orbit of a spacecraft in three dimensional Inertial Reference Frame is in general represented by a standard set of six parameters like Keplerian Orbital Elements namely semimajor axis, eccentricity, inclination, argument of perigee, right ascension of ascending node, and true anomaly. An orbit can also be represented by an equivalent set of six parameters namely the position and velocity vectors, hereafter referred as orbit-vectors. The process of determining the six orbital parameters from redundant set of observations (more than the required minimum observations) is known as Orbit Determination (OD) process. This is, in general, solved using Least Squares principle. Availability of accurate, almost continuous, space borne observations provide tremendous scope for simplifications and new directions in Autonomous OD (AOD). The objective of this thesis is to develop a suitable scheme for onboard autonomy in OD, specifically for low-earth-orbit-missions that are in high demand in the immediate future. The focus is on adopting a simple orbit model by a thorough study and analysis by considering the individual contributions from the different force models or component accelerations acting on the spacecraft. Second step in this work is to address the application of an onboard estimation scheme like Kalman Filter for onboard processing. The impact of the approximation made in the orbit model for filter implementation manifests as propagation error or estimation residuals in the estimation. The normal procedure of tuning the filter is by getting an appropriate state and measurement noise covariance matrices by some means, sometimes through trial and error basis. Since this tuning is laborious and the performance may vary with different contexts, it is attempted to propose a scheme on a more general footing, with dynamically compensating for the model simplification. There are three parts of this problem namely (i) Analysis of different Orbit Dynamics Models and selection of a simplified Onboard Model (ii) Design of an Estimator Filter based on Kalman Filter approach for Onboard Applications and (iii) Development of a suitable Filter Compensation procedure to ensure best estimates of orbit vectors even with the simplified orbit model. Development of a Numerical Integration scheme (and a software tool) and extensive simulation exercises to justify the conclusion on the simple model to be used in the estimation procedure forms the first part of the thesis. Tables quantify the effect of individual accelerations and demonstrate the effects of various model components on orbit propagation. In general, it is well known that the atmospheric drag is a non-conservative force and reduces energy; it is also known that the effect of first zonal harmonic term is predominant than any other gravity parameters; such anticipated trends in the accuracies are obtained. This particular exercise is carried out for orbits of different altitudes and different inclinations. The analysis facilitates conclusions on a limited model orbit dynamics suitable for onboard OD. Procedures and results of this model selection analysis is published in Journal of Spacecraft Technology, Vol. 16, No.1,pp 8-30, Jan 2006, titled “Orbit Model Studies for Onboard Orbit Estimation” [69]. Design of Estimator based on Kalman Filter There are two steps involved in dealing with the next part of the defined work: • Design and implementation of Extended Kalman Filter Estimation (EKF) scheme • Steps to compensate for approximation made in the reduced orbit dynamics The GPS receivers on board some of the IRS satellites (for example, the Resource-Sat-1), output the GPS Navigation Solutions (GPSNS) namely the position and velocity vectors of the IRS satellite along with the Pseudo-range measurements. These are recorded onboard for about two orbits duration, and are down loaded. An Extended Kalman Filter Algorithm for the estimation of the orbit vectors using these GPSNS observations is developed. Estimation is carried out assuming a Gaussian white noise models for the state and observation noises. The results show a strong dependence on the initial covariance of the noise involved; reconstruction of the observations results only if the assumption of realistic noise characteristics (which are unknown) is strictly adhered. Hence this simple non-adaptive EKF is found inadequate for onboard OD scheme. Development of the Dynamics Filter Compensation (DFC) Scheme In next part of the thesis, the problem of dealing with the un-modeled accelerations has been addressed. A suitable model-compensation scheme that was first developed by D.S Ingram el at [60] and successfully applied to Lunar missions, has been modified suitably to treat the problem posed by the reduced orbit dynamics. Here, the un-modeled accelerations are approximated by the OU stochastic process described as the solution of the Langavin stochastic differential equation. A filter scheme is designed where the coefficients of the un- modeled acceleration components are also estimated along with the system state yielding a better solution. Further augmentation to the filter include a standard Adaptive Measurement Noise covariance update; results are substantiated with actual data of IRS-P6 (Resource–Sat 1, see chapter 4). Classified as the Structured Adaptive Filtering Scheme, this results in a Dynamic Filter Compensation(DFC) Scheme which provides distinctly improved results in the position of the state. First, the estimation is carried out using actual GPS Navigation Solutions as observations. What is to be estimated itself is observed; the State-Observation relation is simple. The results are seen to improve the orbit position five times; bringing down the position error from 40 meters to about 8 meters. However, this scheme superimposes an extra factor of noise in the velocity vector of the GPSNS solutions. It is noted that this scheme deals only with the process noise covariance. To tackle the noise introduced in the velocity components, modifications of the original scheme by introducing an adaptive measurement noise covariance update is done. This improves the position estimate further by about 2 meters and also removes the noise introduced in the velocity components and reconstructs the orbit velocity vector output of the GPSNS. The results are confirmed using one more set of actual data corresponding to a different date. This scheme is shown to be useful for obtaining continuous output –without data gaps- of the GPSNS output. Next, the estimation is carried out taking the actual GPS observations which are the Pseudo Range, Range rate measurements from the visible GPS satellites (visible to the GPS receiver onboard ). Switching over to the required formulation for this situation in the state-measurement relation profile, estimation is carried out. The results are confirmed in this case also. Clear graphs of comparisons with definitive orbital states (considered as actual) versus estimated states show that the model reduction attempted at the first part has been successfully tackled in this method. In this era of space-borne GPS observations, where frequent sampling of the orbiting body is suggestive of reduced orbit models, an attempt for replacement of the conventional treatment of expensive and elaborate OD procedure is proved feasible in this thesis work.