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.

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<p>The main theme covered in this thesis is experimentalstudies of quantum dynamics and coherent control in homonuclearalkali diatomic molecules by ultrafast laser spectroscopy iththe implementation of pump-probe techniques.</p><p>A series of experiments have been performed on the Rb2molecules in a molecular beam as well as in a thermal oven. Thereal-time molecular quantum dynamics of the predissociatingelectronically excited D(3)¹Πu state of Rb₂, which couples to/intersects several otherneighbouring states, is investigated using wavepackets. Thepredissociation of the D state, explored by this wavepacketmethod, arises from two independent states, the (4)³Σ_u⁺and (1)³∆_u, for which the second corresponds to a much fasterdecay channel above a sharp energy threshold around 430 nm. Thelifetime of the D state above the energy threshold is obtained,τ ≈ 5 ps, by measuring the decay time of thewavepacket in a thermal oven. Further experimentalinvestigation performed in a molecular beam together withquantum calculations of wavepacket dynamics on the D state haveexplored new probe channels of wavepacket evolution: theD′(3)1Σu+ channel, which exhibits vibrational motionin a shelf state and the (4)³Σu+ channel, where direct build-up of thewavefunction is observed due to its spin-orbit oupling to the Dstate.</p><p>The real-time quantum dynamics of wavepackets confined totwo bound states, A¹Σ_u⁺(0_u⁺) and b³Π_u(0_u⁺), have been studied by experiment andcalculations. It is shown that these two states are fullycoupled by spin-orbit interaction, characterised by itsintermediate strength. The intermediate character of thedynamics is established by complicated wavepacket oscillationatterns and a value of 75 cm^-1is estimated for the coupling strength at thestate crossing.</p><p>The experiments on the Li₂molecule are performed by coherent control ofrovibrational molecular wavepackets. First, the Deutsch-Jozsaalgorithm is experimentally demonstrated for three-qubitfunctions using a pure coherent superposition of Li₂rovibrational eigenstates. The function’scharacter, either constant or balanced, is evaluated by firstimprinting the function, using a phase-tailored femtosecond(fs) pulse, on a coherent superposition of the molecularstates, and then projecting the superposition onto an ionicfinal state using a second fs pulse at a specific delay time.Furthermore, an amplitude-tailored fs pulse is used to exciteselected rovibrational eigenstates and collision induceddephasing of the wavepacket signal, due to Li₂-Ar collisions, is studied experimentally. Theintensities of quantum beats decaying with the delay time aremeasured under various pressures and the collisional crosssections are calculated for each well-defined rovibrationalquantum beat, which set the upper limitsfor ure dephasingcross sections.</p><p><b>Keywords:</b>Ultrafast laser spectroscopy, pump-probetechnique, predissociation, wavepacket, pin-orbit interaction,coherent control, (pure) dephasing</p>

<p>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. </p><p>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.</p>

Some of the outflowing ions in the plasma edge have sufficient energy to access orbits which allow them to free-stream out of the confined plasma region and be lost to the wall or divertor. The effects of this ion-orbit-loss (IOL) on the poloidal distribution of ion, energy and momentum fluxes from the plasma edge into the tokamak scrape-off layer (SOL) are analyzed for a representative DIII-D H-mode discharge. IOL yields large fluxes of particle, energy and momentum, distributed poloidally over the SOL, but predominantly into the outboard SOL, significantly changing the fluxes due to transport processes for confined ions within the edge plasma. An intrinsic co-current rotation in the edge of the plasma is produced by the preferential loss of counter-current ions

One of the most important aspects of preliminary interplanetary mission planning entails designing a trajectory that delivers a spacecraft to the required destinations and accomplishes all the objectives. The design tool described in this thesis allows an investigator to explore various interplanetary trajectories quickly and easily. The design tool employs the patched conic method to determine heliocentric and planetocentric trajectory information. An existing Lambert Targeting routine and other common algorithms are utilized in conjunction with the design tool’s specialized code to formulate an entire trajectory from Earth departure to arrival at the destination. The tool includes many options for the investigator to accurately configure the desired trajectory, including planetary gravity assists, deep space maneuvers, and various departure and arrival conditions. The trajectory design tool is coded in MATLAB, which provides access to three dimensional plotting options and user adaptability. The design tool also incorporates powerful MATLAB optimization functions that adjust trajectory characteristics to find a configuration that yields the minimum spacecraft propellant in the form of change in velocity. / text

The Gravity Recovery and Climate Experiment (GRACE) has been orbiting the Earth and determining its gravity field since 2002. Throughout the course of the mission, the orbital elements occasionally change such that the satellites enter a repeat ground track configuration. Repeat ground tracks result in reduced spatial resolution of the satellites, which poses problems in the context of gravity recovery. The monthly gravity solutions during these periods are examined and shown to have lower quality than usual. The characteristics of these repeat period solutions are identified and compared to a period of uniform coverage to illustrate the ways in which the solutions are degraded. An investigation into the underlying physical and computational sources of these errors is also presented. / text

The laser geodetic satellites Lageos and Starlette exhibit residual orbital motion with an unexplained seasonal component. In addition, recognized polar motion excitation sources do not account for a large portion of observed polar motion. It is hypothesized that air and ocean mass redistribution is the primary source of seasonal perturbations in satellite motion, and that wind-driven ocean mass redistribution is a major source for polar motion excitation. Average monthly variations in zonal spherical harmonic geopotential coefficients are estimated from NMC air pressure for 1958 through 1973, and from variations in continental water storage predicted by a global hydrologic model. These coefficients are used to predict average monthly perturbations in the longitude of the ascending node ([Omega]) for Lageos and Starlette, and in the eccentricity vector ([Psi]) for Starlette. WMO monthly air pressures and twice-daily Navy sea level pressures are used to predict time series of [Omega] and [Psi] perturbations for Lageos during 1976 through 1985, and for Starlette during 1980 through 1983. In addition, the Hellerman and Rosenstein wind stress field for world oceans and the Gill-Niiler bottom pressure equation are used to estimate annual and semi-annual ocean mass redistribution, and to predict polar motion excitation vectors and Lageos [Omega] perturbations. Comparison of predicted [Omega] and [Psi] perturbations with observed Lageos and Starlette behavior indicate that air pressure may be responsible for much of the unmodeled seasonal variation in the Earth's geopotential. In contrast, the water storage contribution is very small. Year-to-year variability in the observed Lageos and Starlette [Omega] times series is well matched by predicted perturbations. Even after the removal of annual and semi-annual components, significant coherence remains between predicted and observed [Omega] time series for both Lageos and Starlette at periods of less than one year. Comparison of predicted polar motion with ILS observations suggest that the effect of ocean mass redistribution is significant, and second only to air pressure in magnitude. Lageos [Omega] perturbations predicted from ocean mass redistribution indicate that non-isostatic sea level fluctuations should be readily observable by satellite laser ranging. / text

A mission architecture consisting of two NASA LONESTAR-2 satellites in Low Earth Orbit is considered. The craft are equipped with cross-communication radios and GPS units. Analyses are conducted for ejection, thruster and attitude maneuvers to achieve objectives of the mission, including sustained communications between the craft. Simulations are conducted to determine the duration of the communication window following the initial separation of the two craft. Recommendations are made to maximize this window while accounting for attitude constraints and the effects of atmospheric drag. Orbital mechanics and control theory are employed to form an algorithm for filtering GPS position fixes. The orbit-determination algorithm accounts for the effects of drag and Earth’s oblateness. Procedures are formed for verifying the initial separation velocities of two spacecraft and for measuring the velocity imparted by impulsive thruster maneuvers. An algorithm is also created to plan the timing and magnitude of corrective thruster maneuvers to align the orbital planes of the two craft. When the craft pass out of communication range, a ground station is used to relay data and commands to conduct state rendezvous procedures. A plan for coordinated attitude maneuvers is developed to strategically utilize the cumulative effects of drag and orbit decay to align the craft over long time periods. The methodologies developed here extend prior research into close proximity operations, forming the foundation for autonomous on-orbit rendezvous under a broader set of initial conditions. / text

This dissertation describes the theoretic studies of magnetic moment and spinorbit interaction in vacuum (Dirac wavepacket) and solid state systems, such as semiconductors. The semiclassical approach developed here provides a simple and intuitive picture for the origin of spin and spin-orbit coupling. In the Dirac model, the spin magnetic moment is originated from the self-rotating Dirac wavepacket with a correct g-value. The spin-orbit interaction is related to Berry connection (gauge potential) and the model is generalized to solid state systems. The Rashba effect caused by the spin-orbit coupling in a crystal with asymmetric potential in heterostructure quantum well is calculated by semiclassical spindependent scattering. The exact treatment of interface phase accumulation provides a justification of spin-dependent boundary condition at interface derived in previous treatment using Löwdin decomposition. Other spin-orbit coupling related phenomena in solid state system are also discussed in this thesis. / text