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The robust loop-shaping control methodology is applied in the flight control system design of the Cranfield A3 Observer unmanned, unstable, catapult launched air vehicle. Detailed linear models for the full operational flight envelope of the air vehicle are developed. The nominal and worst-case models are determined using the v-gap metric. The effect of neglecting subsystems such as actuators and/or computation delays on modelling uncertainty is determined using the v-gap metric and shown to be significant. Detailed designs for the longitudinal, lateral, and the combined full dynamics TDF controllers were carried out. The Hanus command signal conditioning technique is also implemented to overcome actuator saturation and windup. The robust control system is then successfully evaluated in the high fidelity 6DOF non-linear simulation to assess its capability of launch stabilization in extreme cross-wind conditions, control effectiveness in climb, and navigation precision through the prescribed 3D flight path in level cruise. Robust performance and stability of the single-point non-scheduled control law is also demonstrated throughout the full operational flight envelope the air vehicle is capable of and for all flight phases and beyond, to severe launch conditions, such as 33knots crosswind and exaggerated CG shifts. The robust TDF control law is finally compared with the classical PMC law where the actual number of variables to be manipulated manually in the design process are shown to be much less, due to the scheduling process elimination, although the size of the final controller was much higher. The robust control law performance superiority is demonstrated in the non-linear simulation for the full flight envelope and in extreme flight conditions.

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.

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.

Unmanned Vehicles (UVs) currently perform a variety of tasks critical to a military mission. In future, they are envisioned to have the ability to accomplish a mission co-operatively and effectively with limited fuel onboard. In particular, they must search for targets, classify the potential targets detected, attack the classified targets and perform an assessment of the damage done to the targets. In some cases, UVs are themselves munitions. The targets considered in this thesis are stationary. The problem considered in this thesis, referred to as the UV problem, is the allotment of tasks to each UV along with the sequence in which they must be performed so that a maximum number of tasks are accomplished collectively. The maneuverability constraints on the UV are accounted for by treating them as Dubin's vehicles. Since the UVs considered are disposable with life spans governed by their fuel capacity, it is imperative to use their life as efficiently as possible. Thus, we need to develop a fuel-optimal (equivalently, distance optimal) motion plan for the collection of UVs. As the number of tasks to be performed and the number of vehicles performing these tasks grow, the number of ways in which the set of tasks can be distributed among the UVs increases combinatorially. The tasks a UV is required to perform are also subject to timing constraints. A UV cannot perform certain tasks before completing others. We consider a simplified version of the UV problem and do not take into account the timing constraints on the tasks to be performed on targets. We use linear programming and graph theory to find a solution to this simplified UV problem; in the graph theory approach, we develop an algorithm which is a generalization of the solution procedures available to solve the Traveling Salesman Problem (TSP). We provide an example UV problem illustrating the solution procedure developed in this thesis.

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