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An analytical approach for satellite orbit determination methodology will be investigated in this work. The motivation for this study is the realisation that enhanced microsatellites can be expected which require accurate orbital knowledge and control onboard to support their payloads. However, a satellite tracking capability has not yet been implemented at Surrey even on the ground segment, as this would involve a considerable cost to maintain and operate. Advanced modern technology enables micro or nanosatellites to have low cost and low power Global Positioning System (GPS) receivers, which opens the way for onboard orbit determination. However, the heavy computational demand required for executing orbit determination has made such an approach unsuitable given the onboard processing environment. In order to overcome this problem, a novel analytical description of a perturbed orbit has been developed by focusing on near circular orbits which is appropriate for Low Earth Orbit (LEO) satellites. To use this analytic description of the orbit for orbit determination extensively reduces the computational demand as no numerical integrator is required to propagate the orbit, the variational equation of the orbit, or state transition matrix. Thus analytical solutions to the orbit perturbation problem are the main focus of this work. The majority of extensive studies made mainly in the late 50's to 60's on this subject considers the general case of orbits, so that the solutions are extremely lengthy and complex. Due to the conventional methods they have chosen, the solutions have a singularity when the eccentricity approaches zero and the argument of perigee becomes undefined. The new description of the perturbed motion of near circular orbits is developed in this work, which does not have a singularity when the eccentricity is zero. This novel description is a natural expansion of the epicyclic motion of a small eccentric orbit, which results in a much simpler expression in the solutions and provides a greater understanding of orbital geometry. A fully analytical orbit determination system is also developed based on this epicyclic description of a perturbed orbit. As this orbit estimator requires no numerical integration scheme, it has extensive computational advantages and suits most onboard applications. This analytical orbit estimator is practically implemented onboard UoSat-12, which is Surrey's 350 kg class minisatellite. It has been operated for more than two years, and has demonstrated reliable autonomous onboard orbit determination. The in-orbit orbit determination results as well as the evaluations of practical accuracy are also presented.

A novel strategy for automatic satellite constellation design with satellite diversity is proposed. The automatic satellite constellation design means some parameters of satellite constellation design can be determined simultaneously. The total number of satellites, the altitude of satellite, the angle between planes, the angle shift between satellites and the inclination angle are considered for automatic satellite constellation design. Satellite constellation design is modelled using a multiobjective genetic algorithm. This method is applied to LEO, MEO and hybrid constellations. The advantage of this algorithm is automatic satellite constellation design whilst achieving dual satellite diversity statistics. Furthermore a new strategy of dynamic channel allocation is proposed using a genetic algorithm for use in MSS networks. The main idea behind this algorithm is to use minimum cost as a metric to provide optimum channel solutions for specified interference constraints. The frequency reuse condition for all spotbeams is investigated as a function of time. The update interval time and the sampling time are introduced in order to track time valiant coefficients and constraints of the algorithm. The method is demonstrated for S-UMTS based on a MEO satellite constellation. Using this algorithm, it is shown that the proposed model outperforms conventional DCA schemes in terms of capacity of the system and Quality of Service (QoS).We show in the thesis that the genetic algorithm is a robust method for calculation of dynamic variations in satellite constellation design and provides resource allocation improvements over DCA in MSS system networks.

The work presented in this thesis, which details the development and evaluation of an optical fibre damage sensor, was supported by The European Space Administration and by a DTI-NERC LINK grant under the SEASENSE programme. The sensing system developed as part of this programme uses fibre Bragg gratings as sensing elements. The translation of the output of these sensors into strain information is achieved using a novel interrogation system described in this work. The work presented in this thesis is believed to be the first evaluation of this technology for the detection / characterisation of hypervelocity impacts and impact damage. In order to evaluate the suitability of this system for spacecraft impact damage assessment, this work explores the response of aerospace composites to impact by reviewing previous work in composite modelling, acoustics and experimental evaluation. This knowledge is extended by the use of the system in both high and low velocity impact experiments with sensors embedded in carbon fibre reinforced plastic. Alternative optical techniques and conventional sensing methods are compared to the observed performance. The system has also proved to be of sufficient resolution and close to sufficient accuracy for use in monitoring an array of Bragg grating based temperature and pressure sensors for oceanographic applications. This work enabled a detailed examination of the system's accuracy, and stability, which was required for structural sensing studies, and illustrates the breadth of possible applications for this technology. The work concludes that this system would be capable of measuring impacts large enough to cause damage that would threaten the integrity of the composite component. Very few of the large number of small impacts expected would occur close enough to a Bragg grating sensor to register a reading. However, with further development, an interrogation system could be created that was capable of monitoring the impact induced flexural wave in the composite. An array of sensors would allow triangulation of the location of impact, and measurement of the impact: momentum.

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.

An increasing number of multiple-satellite constellations, providing global telecommunications, will be launched into low Earth orbit (LEO) within the next decade. These systems could be utilised for many years in a growing debris environment that presents, a significant long-term collision hazard. The main objective of this PhD is to examine the long-term impact of these systems on the LEO debris environment, and vice versa. The high-resolution simulation of the historical and long-term future evolution of the LEO debris environment, and the detailed long-term prediction of collision risk to any target orbit intersecting LEO, is not trivial to achieve within a single model. The Integrated Debris Evolution Suite (IDES) has been developed as an extremely flexible tool, with a wide scope of state-of-the-art capabilities in all of these areas. A highly novel aspect of the IDES model is the new target-centred approach to the prediction of future collision events. This approach has made an advance over other traditional methods in the area of future collision event prediction and has improved the accuracy of modelling the all-important future collision fragment source within long-term debris evolution models. The IDES simulation software has undergone a rigorous validation programme to assess its accuracy. The historical debris environment simulated by IDES was validated by comparison with reliable measurement data. This validation exercise has greatly improved the confidence in the IDES model for the prediction of the long-term evolution of the debris environment and mission collision risks. The validated IDES model has been used in a number of comprehensive state-of-the-art applications. The long-term collision interactions of a wide range of different constellation designs with the LEO debris environment have been extensively simulated in a 'business as usual' future traffic scenario, both with and without the implementation of debris mitigation measures. The long-term impact of these different constellation designs on the future collision rate, population levels and collision risks in LEO has been evaluated to determine the sizes and orbits of constellations that the LEO debris environment can tolerate. An overall assessment of the long-term impact of the currently foreseen constellation traffic on the LEO debris environment has been performed. The effectiveness of a package of different routine mitigation measures on stabilising the long-term LEO debris environment has been studied and presented in the thesis. Long-term forecasts of LEO constellation collision risk have been produced by the IDES model. These forecasts are state-of-the-art and should be of considerable interest to constellation mission designers. Estimated debris-induced satellite failure rates were assessed for varying constellation architectures in order to find the sensitivity of the predictions to influential design parameters such as orbit selection, the number of satellites in the system, and satellite cross-sectional area. To put these predictions into context, the estimated debris-induced failure rates were compared to the rates expected from satellite component or sub-system failure.

This thesis pioneers the application of nonlinear dynamics to spacecraft power systems. Two areas of general interest are addressed. On the one hand, the fundamental dynamics of space power systems were investigated from a nonlineax dynamics perspective, and on the other, nonlinear dynamics concepts were used to realise a practical engineering application. The former examines four simple but relevant space power system models. The study revealed a variety of bifurcations, coexisting attractors and chaotic behaviour that could potentially shed light on some familiar but poorly understood effects in space power systems operations, including bus voltage collapse, spurious oscillations, and chaotic 'noise'. Because such behaviour manifests itself in nonlinear systems but could not be exposed by customary linear systems theory, potential anomalies may remain unpredicted which could lead to catastrophic consequences. As such, these results have important implications to reliability issues, critical in space. The exposition of the concepts and tools used in this thesis would serve the practising engineer by providing the basis and pave the way for studying larger and more complex systems, in the quest for improved system performance and reliability. In the course of this work, an algorithm to compute the maximum Lyapunov exponent from differential equations with discontinuities was required to confirm chaos. Although the concepts and tools for investigating smooth equations are well established, dynamics of non-smooth systems have not been extensively studied. Here, the algorithm proposed by Miiller to cope with the discontinuities in mechanics was reviewed and was found to be applicable to power electronics in general. As a confirmation, this algorithm was applied successfully to a well known Buck DC-DC converter. Although the exploitation of nonlinear dynamics to engineer direct practical applications is still in its infancy, one is presented in this thesis. A maximum power point tracker was synthesised via nonlinear dynamics principles, simulated and experimentally verified. Excellent static and dynamic performance were exhibited. In addition, a two-dimensional stroboscopic map was derived which adequately described the fundamental dynamics of the system. This is confirmed from the good agreement between the simulated and experimental return maps. Via this map and further bifurcation study, preliminary design guidelines were established.

This thesis presents an investigation into methods for predicting and improving the reliability of electronic subsystems on low cost small satellites [LCSS]. Current methods for predicting the reliability of electronic systems were identified and the feasibility of these methods for predicting the reliability of electronic systems on LCSS were analysed. Trends in satellite failures and satellite design methodologies were studied. It was identified that, highly capable, LCSS have became possible due to the existence of advanced commercial-off-the-shelf (COTS) very large scale integrated circuits (VLSI) components. Typically, a large number of these components are employed on these satellites and the effect of space radiation on these components is a significant source of risk to their reliability. This research program therefore concentrated on addressing the effects of space radiation on the reliability of COTS VLSI components. In particular, traditional methods for predicting the rate and effects of radiation induced spontaneous logic state changes (Single Event Upsets (SEUs)) in satellite on-board microprocessors were critically examined. Problems associated with these methods were identified, and two new models to enable more accurate predictions were proposed, developed and tested. One of the models, a duty cycle prediction model, enables a better prediction of the "observable SEU induced error rate" in satellite microprocessor systems; and the other model, a SEU simulation model, enables the nature of SEU induced errors observed at microprocessor system level to be studied. The practical applications of both models were demonstrated by analysing a real satellite application system. These models will be particularly useful to engineers in the selection and application of reliable COTS microprocessors for satellite onboard computer intensive tasks without jeopardising reliability.

Control of the inter-satellite distances within a cluster of colocated satellites located in the same GEO window is examined with regards to the close approaches between pairs of satellites. Firstly, the orbital evolution and station keeping control of a single GEO satellite is examined and a new IBM PC based software program capable of performing both these functions autonomously from initial values of the orbital position and date is detailed and validated. Cluster design ideas are then examined in detail and the propagation software is used to generate data for a cluster of four satellites. Two test cases are examined to quantify the frequency of close approaches between individual satellite pairs, each test case using a different orbital element separation strategy but the same station keeping control scheme. The results of the study are then compared with previous research and discussions are presented on the advantages of each method. Finally, a cluster geometry correction manoeuvre, based on Hill's equations of relative motion, is presented which requires only those thrusters used by typical station keeping. This manoeuvre is integrated into the computer software and the two test cases noted previously are again propagated and the close approach results analysed to demonstrate the reduction in the number of close approaches below 5 km.

This thesis describes the definition, implementation, and in-orbit testing of an autonomous navigation unit based upon a GPS receiver for use on board a small satellite in low Earth orbit. It explains the motivation for the use of GPS to provide this function, and describes the practical application and integration of this technology into an existing microsatellite system. Until now, the technology for any satellite to track itself has not existed. Space agencies spend significant funds supporting a network of tracking stations around the world for orbit determination. With the recent realisation of the Global Positioning System and the availability of inexpensive receiver hardware, it has become a practical proposition to include a GPS receiver within the demanding constraints of a small satellite. A GPS receiver on-board a satellite can eliminate the necessity for ground-based tracking by providing an autonomous orbit determination capability. During the course of these studies, the requirements and constraints of a small satellite were identified by the author and matched with the capabilities of a GPS receiver. A GPS Navigation Unit was defined to provide autonomous services available oil demand for the satellite platform and payloads; position and velocity; time synchronisation; orbital elements; payload triggering and GPS data logging (for experimental and research purposes). The GPS Navigation Unit includes a processing facility capable of command and initialisation of the GPS receiver, and data processing to give orbit determination capability. When used on a microsatellite, the additional constraints of low power consumption necessitate the intermittent operation of the GPS receiver. To test the concept of the GPS Navigation Unit, a commercial Trimble TANS II GPS receiver system that had been modified for orbital velocities was integrated into the PoSAT-1 microsatellite which was launched into low Earth orbit in September 1993. A method for orbit determination was developed for use with the output from the GPS receiver, and the GPS Navigation Unit was implemented in software according to the constraints of the PoSAT-1 mission. The significant results from these studies include: The first use of a GPS receiver on a microsatellite, PoSAT-1. The implementation, test and validation of a GPS Navigation Unit in low Earth orbit. The first satellite mission to demonstrate the capability for autonomous orbit determination through the GPS Navigation Unit. The definition of the general-purpose interfaces between a small satellite and a satellite- borne GPS Navigation Unit.

The establishment of large space systems in the Earth's orbit requires the development of new construction techniques. As large structures are being proposed for future projects, several orbit construction processes have evolved. Of these, the concept of deployable truss structures is one which is appraised during this research. An energy loaded joint is presented which realises the potential of this technique. Incorporated into a skeletal truss framework, however, the structural characteristics of the joint will influence the overall performance of the structure. In order to assess the significance of this, an investigation is conducted into the structural behaviour of the joint. Although the deployable concept has been primarily developed for realising large structures in the Earth's orbit, the technique can also be used in terrestrial applications. Accordingly, a numerical study evaluates the static stress distribution generated through the joint; the structure is considered to be in a fully deployed state. The investigation appraises the effects of manufacturing tolerances and indicates their significance on design. Dynamic loading induced in the space environment highlights the joint to possess a strong geometrical nonlinearity. The joint describes a bilinear flexural stiffness property that is associated with the design. As a consequence, resonant frequencies exhibited by small truss structures are found to be dependent on a variety of geometric and loading parameters. Large trusses are also found to be affected although their behaviour is seen to be less dependent on the characteristics of individual joints. An experimental investigation confirms the presence of nonlinearity within the energy loaded joint. The characteristic response during vibration highlighted the difficult task of predicting structural behaviour. This instigated a redevelopment of the joint design incorporating measures which eradicate the cause of the geometrical nonlinearity. Tests reveal this design to be a significant improvement and an important contribution to the development of deployable structures.

The demand for the mobile communications is growing, as is the trend for the introduction of new services and for these to be available over wide geographical areas. Satellites, due to their inherent ability to support long distance reliable communication, are believed to play a pivotal role in realising such systems at national and international levels. Until the recent past, satellites were used to provide a communication facility to mobile platforms that could afford to carry relatively heavy and complex terminals such as those in maritime use. However, many mobile services for aeronautical and land users have also been introduced since 1990. Land Mobile Satellite (LMS) communications have until now been restricted to low data rates and geostationary orbit (GEO) satellites. Increasingly the demand is for integrated speech and data services. Signal propagation in land mobile conditions is one of the most important design considerations limiting the performance of a satellite based mobile communication system. In order to provide communication services through a viable satellite network, a proper knowledge of link degradation due to the mobile channel conditions is essential. In the land mobile environment, shadowing of the line-of-sight (LOS) signal and multipath propagation can limit the link performance for considerable percentages of time. Satellites in orbits other than the geostationary orbit have, therefore, been proposed for mid-to-high latitude regions in order to provide high elevation angles at the mobiles, thus reducing losses due to blockage, and as a result improve link margin requirements. In addition, to achieve world-wide mobile communications, constellations of low earth orbit (LEO) satellites have been proposed. However, sufficient information on link performance, under reduced shadowing conditions at high elevation angles, does not exist and this is critically important in the planning or implementation of future HEO and LEO based satellite systems. The research work reported here was carried out in an attempt to extend the existing propagation database applicable to high elevation angle land mobile satellite (LMS) channel at the L-band. However, other frequencies corresponding to the S and Ku bands, were also included to generate data at frequencies for likely future applications and enable a direct comparison of propagation characteristics to be made between the bands. In the absence of appropriate satellite source(s), the multiband propagation measurement campaign was undertaken using a helicopter. The helicopter provided a platform which simulated the high elevation angle satellite signal. Different environments such as suburban, wooded and open were considered, and mobile measurements were made at various high path elevation angles in each environment. The statistical analysis of the measurement data provided useful information on LMS channel behaviour and its dependency on the signal frequency, elevation angle and the mobile channel environment. It has been observed that smaller attenuations are experienced as the elevation angle increases or the radio frequency decreases. Significant effect on the channel fading is observed due to variable amounts of shadowing encountered under varying channel environments. Using the results of the measurement campaign described in this thesis, it has been possible to propose a fading model for use in future planning for high elevation paths that will be encountered in both HEO and LEO mobile satellite systems.

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.

An important concern in spacecraft engineering is attitude determination control and stabilisation (ADCS) - the combination of applied mathematics, classical physics and modern technology which maintains the pointing direction of one or more axes of an Earth-orbiting satellite. This thesis is a detailed study of a particular type of ADCS which exploits the gravity-gradient effect, which is just the weak tendency for an appropriately shaped body to point naturally in preferred directions; reinforced by magnetorquing, which is the active interaction of the geomagnetic field with a switched current passing through coils in the spacecraft body. The advantages of this technology is that it is low-cost, non-consumable and has no moving parts - so constituting no limitation to the satellite's life. The thesis is a detailed study of this form of ADCS with specific application to low Earth polar-orbiting (LEO) satellites, for which it is particularly suitable. The work is also a study in attitude determination based solely on a 3-axis magneometer measurement of the geomagnetic field, which is in principle a simple way, in terms of technology, of determining the attitude of the spacecraft, and from this controlling the attitude, if mediated by an on-board computer implementing appropriate algorithms. The results are for the most part practically based on the author's involvement with two satellites over a six year period with the satellites UOSAT-1 and UOSAT-2, which were designed, built, and continue to be controlled from the University of Surrey. A practical innovation in 2-axis attitude control is described: the active 'delibration' by active damping of a gravity-gradient controlled LEO satellite in an attitude-stabilised state, using a threshold comparison algorithm. A new theory and algorithms are then developed for 3-axis attitude control, based on a complementary use of magnetorquing and gravity gradient. Subject to further development these could alter the general perception of the most effective way of controlling low Earth orbiting satellites.