A perfect duality exists between the problem of space-based orbit determination from line-of-sight measurements and the problem of designing an interplanetary autonomous navigation system. Mathematically, these two problems are equivalent. Any method solving the first problem can be used to solve the second one and, vice versa. While the first problem estimates the observed unknown object orbit using the known observer orbit, the second problem does exactly the opposite (e.g. the spacecraft observes a known visible planet). However, in an interplanetary navigation problem, in addition to the measurement noise, the following "perturbations" must be considered: 1) light-time effect due to the finite speed of light and large distances between the observer and planets, and 2) light aberration including special relativistic effect. These two effects require corrections of the initial orbit estimation problems. Because of the duality problem of space-based orbit determination, several new techniques of angles-only Initial Orbit Determination (IOD) are here developed which are capable of using multiple observations and provide higher orbit estimation accuracy and also they are not suffering from some of the limitations associated with the classical and some newly developed methods of initial orbit determination. Using multiple observations make these techniques suitable for the coplanar orbit determination problems which are the case for the spacecraft navigation using visible planets as the solar system planets are all almost coplanar. Four new IOD techniques were developed and Laplace method was modified. For the autonomous navigation purpose, Extended Kalman Filter (EKF) is employed. The output of the IOD algorithm is then used as the initial condition to extended Kalman filter. The two "perturbations" caused by light-time effect and stellar aberration including special relativistic effect also need to be taken into consideration and corrections should be implemented into the extended Kalman filter scheme for the autonomous spacecraft navigation problem.

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A satellite in orbit about a planet needs some means of attitude control in order to, for instance, get as much sun into its solar-panels as possible. It is easy to understand that, for example, a spy satellite has to point at a certain direction without the slightest trembling to get a photo of a certain point on the earth. This type of mission must not exceed an error in attitude of more then about 1/3600 degrees. But, since high accuracy equals high cost, it is also easy to understand why a research satellite measuring solar particles (or radiation) in space does not need high accuracy at all. A research vessel of this sort can probably do with less accuracy then 1 degree. The first part of this report tries to explain some major aspects of satellite space-flight. It continues to focus on the market for small satellites, i.e. satellites weighing less than 500 kg. The second part of this final thesis work deals with the development of a program that simulates the movement of a satellite about a large celestial body. The program, called AOSP, consists of user-definable packages. Sensors and estimation filters are used to predict the satellites current position, velocity, attitude and angular velocity. The purpose of the program, which is written in MATLAB, is to easily determine the pointing accuracy of a satellite when using different sensors and actuators.

Adaptive antenna arrays have a great importance in reduction of the effect of interference and increase the capacity for the mobile satellite communication. Interference and multipath fading remain a main problem for reception of signals. These two problems obviously affect the overall capacity.  Adaptive antenna arrays in the handheld mobile apparatus will be the solution for the above two problems.   Satellite mobile communication is one of the growing fields in the communication area where terrestrial infrastructures are unable or ineffective to supply. Maritime, aeronautical and land mobile are some of the applications. During natural disasters where ground services are stopped, mobile satellite communications has great importance. Following the hurricane season, the Asian Tsunami and the devastating Haiti earthquake, mobile satellite communications had played a great role to fill the communication gaps.  The satellites can be tracked automatically by adaptive antenna array when it moves in its orbital plane.   In this thesis the methods that how the adaptive antenna array combats interferers is presented and simulated using MATLAB software. The performance of the adaptive antenna array is evaluated by simulating the directivity pattern of the antenna and Mean Square Error (MSE) graph for different scenario like Signal to Interference Noise ratio (SINR), number of iterations, antenna array elements and convergence factor (μ), assuming the signals are coming from different Direction of Arrival (DOA).

The ability to observe resident space objects (RSOs) is a necessary requirement for space situational awareness. While objects in a Low-Earth Orbit are easily ob- servable by ground-based sensors, diffculties arise when trying to monitor objects with larger orbits far above the Earth's surface, e.g. a Geostationary Orbit. Camera systems mounted on satellites can provide an eff ective way to observe these objects. Using a satellite with a speci c orbit relative to the RSO's orbit, one can passively observe all the objects that share the RSO's orbit over a given time without active maneuvering. An orbit can be defi ned by ve parameters: semi-major axis, eccentricity, right ascension of ascending node, inclination, and argument of perigee (a; e; ; i; !). Using these parameters, one can create an orbit that will surround the target orbit allowing the satellite in the Rock-Around Orbit (RAO) orbit to have a 360 degree view of RSOs in the target orbit. The RAO orbit can be applied to any circular or elliptical target orbit; and for any target orbit, there are many possible RAO orbits. Therefore, diff erent methods are required to narrow down the selection of RAO orbits. These methods use distance limitations, time requirements, orbit perturbations, and other factors to limit the orbit selections. The first step is to determine the range of RAO semi-major axes for any given target orbit by ensuring the RAO orbit does not exceed a prescribed maximum al- lowable distance, dmax from the target orbit. It is then necessary to determine the eccentricity range for each possible RAO semi-major axis. This is done by ensuring the RAO still does not exceed dmax but also ensuring that the RAO orbit travels inside and outside of the target orbit. This comprises one half of the rock-around motion. The final step is to determine the inclination of the RAO orbit. Only a small inclination different from that of the target orbit is required to complete the rock-around motion while the maximum inclination is found by making sure the RAO orbit does not exceed dmax. It is then important to consider orbit perturbations, since they can destroy the synchronization between the RAO and target orbit. By examining the e ffects of the linear J2 perturbations on the right ascension of ascending node and argument of perigee, the correct semi-major axis, eccentricity, and inclination can be chosen to minimize the amount of fuel required for station keeping. The optimal values can be found by finding the Delta v needed for di fferent combinations of the variables and then choosing the values that provide the minimum Delta v. For any target orbit, there are multiple RAO orbit possibilities that can provide 360 degree coverage of a target orbit. Even after eliminating some of them based on the methods already described, there are still many possibilities. The rest of the elimination process would then be based on the mission requirements which could be the range of an on-board sensor, the thruster or reaction wheel controls, or any other number of possibilities.

Nowadays, Global Positioning System (GPS) which provide global coverage, continuous tracking capability and high accuracy has been preferred as the primary tracking system for onboard real-time precision orbit determination of Low Earth Orbiters (LEO). In this work, real-time orbit determination algorithms are established on the basis of extended Kalman, unscented Kalman, regularized particle, extended Kalman particle and extended H-infinity filters. Particularly, particle filters which have not been applied to the real time orbit determination until now are also performed in this study and H-infinity filter is presented using all kinds of real GPS observations. Additionally, performance of unscented Kalman filter using GRAPHIC (Group and Phase Ionospheric Correction) measurements is investigated. To evaluate performances of all algorithms, comparisons are carried out using different types of GPS observations concerning C/A (Coarse/Acquisition) code pseudorange, GRAPHIC and navigation solutions. A software package for real time orbit determination is developed using recursive filters mentioned above. The software is implemented and tested in MATLAB&copy / R2010 programming language environment on the basis of the object oriented programming schema.

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In this thesis, we study the normally elliptic singular perturbation problems including both finite and infinite dimensional cases, which could also be non-autonomous. In particular, we establish the existence and smoothness of O(1) local invariant manifolds and provide various estimates which are independent of small singular parameters. We also use our results on local invariant manifolds to study the persistence of homoclinic solutions under weakly dissipative and conservative per- turbations. We apply Semi-group Theory and Lyapunov-Perron Integral Equations with some careful estimates to handle the O(1) driving force in the system so that we can approximate the full system through some simpler limiting system. In the investigation of homoclinics, a diagonalization procedure and some normal form transformation should be first carried out. Such diagonalization procedure is not trivial at all. We discuss this issue in the appendix. We use Melnikov type analysis to study the weakly dissipative case, while the conservative case is based on some energy methods. As a concrete example, we have shown rigrously the persistence of homoclinic solutions of an elastic pendulum model which may be affected by damping, external forcing and other potential fields.

The word 'rendezvous' implies a pre-arranged meeting between two entities for a specific purpose. This term is used in the study of spacecraft operations, to describe a set of maneuvers performed by two spacecraft in order to achieve a match in position and velocity. The term 'orbital rendezvous' applies to rendezvous between spacecraft in earth-centered orbits. Considering its obvious scope for application in the assembly, maintenance and retrieval of earth satellites, the importance of orbital rendezvous towards maintaining a sustained presence in space can be easily appreciated. This particular study deals with the development of a guidance scheme for an orbital rendezvous operation, wherein only one of the spacecraft, called the chaser, is assumed to be provided with a capability to maneuver, while the other spacecraft, the target, is assumed to be thrust-free or passive. There is presently a lot of interest in autonomous trajectory planning and guidance schemes for orbital rendezvous missions. Autonomy here, refers to the absence of ground supervision and control over the on-board planning and guidance process, and is expected to result in greater mission flexibility and lower operating costs. The terms trajectory planning and guidance collectively refer to the optimization process used to determine minimum-fuel trajectories, and the means employed to make the spacecraft follow them, based on navigational updates. The challenge lies mainly in making the autonomous scheme real-time implementable, and at the same time compatible with the limited computational capabilities available on-board. It is well known that a large part of the computation times and costs, when determining optimal trajectories, are taken up by (1) the prediction of spacecraft motion using numerical integration schemes, and (2) the use of iterative numerical techniques to solve the non-linear, coupled system of equations obtained as boundary conditions in the trajectory optimization problem. There exists on the other hand, a wealth of results from analytical investigations into the motion of spacecraft, that can be profitably utilized by use of suitable assumptions, to reduce computation times and costs relating to trajectory prediction. The present thesis seeks to follow this course, while trying to ensure that the assumptions made do not influence in a negative manner the accuracy of the guidance scheme. The assumptions to be described below are based on the division of the total rendezvous maneuver into sub-phases. The trajectory optimization problems for the individual sub-phases are first considered independent of one another. A method is then found to combine the two sub-phases in an optimal manner. The initial or the homing phase of the rendezvous maneuver, consists of an open-loop orbit transfer, intended to place the chaser within a 'window of proximity' spanning a few hundreds of kilometers, of the target. In order to avoid time consuming numerical integration of the non-homogeneous, non-linear central force-field equations of motion, an impulsive thrust model is assumed. A parametric optimization method is used to determine the location, orientation and magnitude of the impulses for a minimum-fuel rendezvous transfer, as it is well known that parametric optimization methods are robust compared to the more general functional optimization methods. A two-impulse transfer is selected, knowing that at least two-impulses are required for a rendezvous maneuver, and that methods are available if necessary, to obtain optimal multi-impulse trajectories from a two-impulse solution. The total characteristic velocity, a scalar cost function related to fuel-consumption, is minimized with respect to a set of independent variables. The variables chosen in this case to determine the rendezvous transfer are (1) the transfer angle θc defining an initial coast in the chaser orbit C by the chaser, (2) the transfer angle θs defining a coast by the target to the position of the second impulse in the target orbit S and (3) a parameter (say p ) that determines the shape of the transfer orbit T between the first and second impulses.