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To measure the satellite-ocean altitude, a radar altimeter transmits a nadir-directed microwave pulse and times the return of the surface reflection. The intervening free electrons of the ionosphere cause group delay of the pulse resulting in an overestimate of the platform altitude by an amount directly proportional to the sub-satellite electron content. In effect the figure of the ocean surface detected by the altimeter is modulated by the spatial and temporal variation of the ionospheric electron content. A two stage technique has been developed to remove the bias imposed by the ionosphere on altimetric measurements. The first stage generates a prediction of electron content based on ionospheric climatology. The second stage is an adaptive modelling procedure which makes use of data from satellite-ranging radar systems. The first chapter of this thesis gives an introduction to the Earth's ionosphere, describes its effect on radar altimetry and suggests a technique to correct for this influence. Chapter Two reviews previous work in related areas before Chapter Three embarks on a description of the spatial and temporal behaviour of electron content. Chapter Four describes the mathematical sub-models which form the basis of the empirical model and Chapter Five is devoted to the calibration and validation of this model. Chapter Six covers the calculation of the coherence functions of electron content which are crucial for the operation of the adaptive procedure. Chapter Seven compares the new model with one employed for a previous altimeter mission and Chapter Eight summarizes what has gone before and suggests topics for future research.

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The cost-effectives mall spacecrafht as becomea n enablingt ool in the pursuit of near earth space commerce. Although small spacecraft have typically forgone the complexity and historically high cost of spacecraft propulsion, the inability to cost-effectively reach specific data gathering orbits from secondary launches presents a serious limitation to the small spacecraft industry. A cost-effective propulsion system capable of moving the secondary spacecraft from the launch orbit to the required mission orbit will effectively increase the number of viable secondary launch opportunities and in some cases provide a higher scientific or commercial return. Propulsion will also allow the dispersing of multiple spacecraft from a single launch vehicle and the inherent ability to de-orbit after a useful mission life. While other propulsion alternatives were considered in this research program, the hybrid rocket was identified as having high potential for suiting the established high-performance, lowcost and safety criteria. However, as this research has shown, the conventional hybrid rocket is not well suited to incorporation within small spacecraft; this is primarily due to the required length verses diameter (UD) to achieve high performance in the conventional hybrid. This research program has produced and tested a novel hybrid rocket engine. The all-new engine is significantly different from the conventional hybrid, exhibiting higher performance and with a geometry that drastically reduces hybrid rocket integration and operation issues. In addition, the new hybrid design has been successfully tested at higher volumetric loading factors than the conventionadl esignsi dentifiedi n the literature. The new alternative geometry hybrid rocket employs tangential oxidiser injectors that induce a vortex flow field to the centrally mounted rocket nozzle. The induced flow field has been shown to provide better fuel and oxidiser mixing. In addition, the tangential oxidiser injection provides an inherent film cooling effect for the combustion chamber wall, allowing the chamber to be fabricated of low cost materials. The new hybrid rocket engine was dubbed the Vortex Flow "Pancake" hybrid or "VFP". This researchp rogramr epresentsth e most technologicallya mbitiousp ropulsionr esearch program conducted by the Surrey Space Centre to date as the tools to analyse and design this engine had to be experimentally derived. Although the fundamental process of burning solid fuel remains unchanged, the combustion chamber gas-dynamics - so vital for predicting fuel liberation and performance within the conventional hybrid - are radically changed in the new configuration. Whereas the conventional hybrid has demonstrated a strong correlation with increasing combustion port diameter and fuel liberation, this research has shown that fuel liberation within the VFP does not obey any such relationships. Operationally, this research has shown that the VFP exhibits a higher fuel volumetric loading factor, higher combustion efficiency and less of an O/F (and consequent performance) shift than conventional designs. This research has proven the VFP to be superior to the conventional hybrid design in every aspect tested. However, this is only part of the benefit realised by the new VFP design as the external geometry of the VFP is the primary benefit enabling the technology to be applied to small spacecraft. Conventional hybrids need L/D ratios in excess of 15 to provide adequate performance, the novel VFP design has been regularly tested at UD's less than 1 with combustion efficiency very near 100%. This unique hybrid characteristic allows the VFP to be integrated on the outside of a spacecraft, in or as part of the spacecraft separation system. An externally mounted engine conserves centrally located spacecraft volume (reducing the need for multiple oxidiser tank scenarios). In addition, the external mount also allows waste heat to be radiated to space rather than other (internal) spacecraft components.

Formation flying is a new paradigm in space mission design, aimed at replacing large satellites with multiple small satellites. Some of the proposed benefits of formation flying satellites are: (i) Reduced mission costs and (ii) Multi mission capabilities, achieved through the reconfiguration of formations. This dissertation addresses the problems of initiatialization, maintenance and reconfiguration of satellite formations in Earth orbits. Achieving the objectives of maintenance and reconfiguration, with the least amount of fuel is the key to the success of the mission. Therefore, understanding and utilizing the dynamics of relative motion, is of significant importance. The simplest known model for the relative motion between two satellites is described using the Hill-Clohessy-Wiltshire(HCW) equations. The HCW equations offer periodic solutions that are of particular interest to formation flying. However, these solutions may not be realistic. In this dissertation, bounded relative orbit solutions are obtained, for models, more sophisticated than that given by the HCW equations. The effect of the nonlinear terms, eccentricity of the reference orbit, and the oblate Earth perturbation, are analyzed in this dissertation, as a perturbation to the HCW solutions. A methodology is presented to obtain initial conditions for formation establishment that leads to minimal maintenance effort. A controller is required to stabilize the desired relative orbit solutions in the presence of disturbances and against initial condition errors. The tradeoff between stability and fuel optimality has been analyzed for different controllers. An innovative controller which drives the dynamics of relative motion to control-free natural solutions by matching the periods of the two satellites has been developed under the assumption of spherical Earth. A disturbance accommodating controller which significantly brings down the fuel consumption has been designed and implemented on a full fledged oblate Earth simulation. A formation rotation concept is introduced and implemented to homogenize the fuel consumption among different satellites in a formation. To achieve the various mission objectives it is necessary for a formation to reconfigure itself periodically. An analytical impulsive control scheme has been developed for this purpose. This control scheme has the distinct advantage of not requiring extensive online optimization and the cost incurred compares well with the cost incurred by the optimal schemes.

The development of a multi-mission small spacecraft power system is described. This system has been designed for the University of Toronto Space Flight Laboratory Generic Nanosatellite Bus (GNB), an approximately 20cm cubical spacecraft with no deployed solar arrays. The GNB is inherently power-generation limited, and consequently, all available power must be utilized with maximum efficiency. This efficiency is achieved using an unconventional parallel-regulated architecture with Peak Power Tracking (PPT) functionality, and is shown to be the PPT design of highest efficiency for spacecraft of this class. In support of this design, a novel spacecraft power simulation suite has been developed, enabling parametric satellite power analysis with high fidelity. Finally, a unique variation on peak power tracking---referred to as peak current tracking---is described. This method is shown to reduce battery depth-of-discharge by as much as 20% over baseline architectures, and furthermore exhibits beneficial emergent behaviour for battery charge management.

The development of a multi-mission small spacecraft power system is described. This system has been designed for the University of Toronto Space Flight Laboratory Generic Nanosatellite Bus (GNB), an approximately 20cm cubical spacecraft with no deployed solar arrays. The GNB is inherently power-generation limited, and consequently, all available power must be utilized with maximum efficiency. This efficiency is achieved using an unconventional parallel-regulated architecture with Peak Power Tracking (PPT) functionality, and is shown to be the PPT design of highest efficiency for spacecraft of this class. In support of this design, a novel spacecraft power simulation suite has been developed, enabling parametric satellite power analysis with high fidelity. Finally, a unique variation on peak power tracking---referred to as peak current tracking---is described. This method is shown to reduce battery depth-of-discharge by as much as 20% over baseline architectures, and furthermore exhibits beneficial emergent behaviour for battery charge management.