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First generation digital codecs for speech have been the subject of worldwide studies since the sixties. Speech quality very close to that of the standard 64 Kbit/s PGM, the ability to carry voice band data all at reasonable complexity and cost are the main objectives of candidate coding techniques. This work deals with the subjective performance evaluation of two types of reduced bit-rate encoding techniques operating at 32 Kbit/s, for inclusion into business satellite communication systems. They are CVSDM and ADPCM, which are both capable of doubling the existing transponder capacity available in business satellite systems. The aim has been to determine transmission planning figures for their quality in terms of Richards' Q-factor. The work is also involved with the implementation of different ADPCM algorithms using digital signal processor (DSP) single chips which is fairly new to the field of telecommunications. Throughout the work the replacement of 64 Kbit/s PCM coder by 32 Kbit/s coder in satellite communication systems has been the theme. The achievement of acceptable subjective quality being the major objective.

This thesis is a theoretical investigation into the thermal and structural response of large reflector support trusses subject to the space environment. A review of the elements necessary in a successful spacecraft design is included. The determination of the temperature distributions within large orbiting tetrahedral trusses is described. The thermal environment is shown to be an important factor'' in determining the performance of spacecraft. High quality microwave performance of the communication system depends upon the magnitude of distortions within the reflector support truss, and the thermal expansion or contraction of members makes a significant contribution to this distortion. The possibility of flexural vibration of individual members of the truss arising during entry into, or exit from, the earth's shadow is considered. Individual slender members and the radio frequency reflecting mesh cast shadows within the truss. A method for calculating the duration and position of such shadowing events is discussed. The effect that shadows have upon shadowee temperatures is determined using computer programs written by the author. Situations in which temperature gradients between the truss faces exist are highlighted. Such gradients produce bending deflections of the truss. Classical theories are developed to predict the magnitude of thermally induced stresses in the adhesive layer of tubular lap joints having dissimilar adherends. Tubular joints which contain a transition layer are also examined. An appropriately chosen transition layer will reduce the adhesive shear stresses by a substantial amount. Finally, the development of thermally induced stresses and damage mechanisms in laminated composite tubes is discussed. The designer must be aware of the amount of damage that spacecraft materials will experience during the long life missions envisaged for the near future.

The research described in this thesis was carried out in four specific areas, namely, Satellite-Switched Time-Division Multiple-Access (SSTDMA) network synchronisation, frame synchronisers, remote sensing satellite data decommutating systems, and prototype frame synchroniser implementation. In the area of SSTDMA network synchronisation, this thesis reports several new synchronisation methods based on an original on-board synchronisation concept. These new methods are shown to have some significant advantages over existing methods. Also reported is a new algorithm to determine the exact satellite position for the open-loop three ranging stations method. In the area of frame synchronisers, adaptive control strategies and post-detection processing techniques are proposed for a new form of frame synchroniser. This new frame synchroniser is shown to provide reliable and optimum frame synchronisation operation in an unpredictable noisy environment. A high-speed version of the new frame synchroniser is also proposed, and it is shown that the adaptive control strategies may be implemented using a microprocessor-based system. In the area of remote sensing satellite data decommutating systems, this thesis reports a new system. Compared with existing systems, the proposed new system offers greater flexibility and expandability, with data decommutation and distribution carried out in real-time. A prototype adaptive frame synchroniser with post-detection processing was constructed, and tested under various simulated environmental conditions. The design and hardware implementation of the prototype frame synchroniser are described. The abilities of the prototype frame synchroniser, such as various mode transition strategies, automatic polarity correction, bit slippage tolerance up to ±2 bits and a bit error rate (BER) tolerance up to 0.208, were demonstrated via a range of tests, which are described in this thesis.

Designing a spacecraft electrical power system (SEPS) is a complex and time-consuming engineering task that involves meeting several design objectives under constraints. A conceptual design of a spacecraft power system involves an optimal selection of available technologies for various components, such as solar cells, solar arrays, batteries, and bus voltages. Each technology has its own advantages and disadvantages that need to be taken into account in the search for an optimal design solution. This selection must meet certain criteria, the most important of which are cost-effectiveness, mass and performance. Traditionally, this task is a manual iterative process. At present, designs thus selected may not be realizable using the state-of-art design options available in the industry. However, advances in domain knowledge and in extra-numerical and multi-objective search techniques, such as evolutionary computation, offer a possibility of accelerating and improving this design cycle through a machine-automated design procedure. This thesis addresses the key issue of intelligent design automation and optimization of spacecraft power systems implemented in realistic design processes. The SEPS design is multi-objective in nature, a situation where a designer searches for solutions that are feasible with respect to all conflicting objectives. To facilitate the intelligent search process, meta-heuristics techniques are exploited in this work to provide computationally inexpensive design optimization. It extends the existing concept of computer-aided design to computer-automated design. To make the process of trade selection more efficient and reliable, a multi-objective design system for solving preliminary design problems for spacecraft electrical power subsystems is developed. It presents a system engineering framework that places design requirements at the core of the design activities. The thesis presents how simulation and optimization techniques can be used to automate and improve the design process of spacecraft power subsystems. The automated design procedure involves the design parameterization and the tools for system sizing and analysis. For the SEPS analysis, an inexpensive method for estimating design behavior is presented. Truly multi-objective and globally optimal design solutions are then artificially evolved as a result of interfacing evolutionary computation techniques with system sizing and analysis tools under practical constraints. Compared with conventional optimization techniques, the multi-objective design approach provides system designers with a clearer understanding of the effect of their design selections on all design variables simultaneously. In particular, the thesis extends a SEPS design problem from the basic technology selection to a detailed optimization based systematic design, which ensures the optimality and usability of designs from the beginning of the design process. Designs are made with implementation of solar cell modeling and parameter optimization using simulated annealing, which forms a very useful tool for simulating the behavior of solar arrays comprising of different types of solar cells. SEPS simulation is extended in MATLAB from existing work currently limited to Si solar cells and NiH2 batteries to a variety of solar cell and battery technologies. The thesis also develops a complete SEPS design and search framework, as a single tool and thus avoiding all compatibility issues involved. This feature makes this work very practical and efficient. It also keeps a way open for further improvements and modifications, both for optimization techniques and for the SEPS search space

The current method of operating space-based assets involves the design and launch of large, monolithic spacecraft. These spacecraft are not responsive to failures or changes in mission requirements, as both require the launch of a completely new spacecraft. The concept of fractionated spacecraft was introduced in 2006 by the US Defence Advanced Research Projects Agency (DARPA) as a way of designing and operating space systems which is more responsive than current methods. The fractionated concept involves the decomposition of the traditional monolithic satellite into a number of free-flying spacecraft connected wirelessly. The free-flying satellites each carry a subset of the required subsystems and share their fractionated resources to achieve the mission objectives. Since 2006, studies of this concept have focused on analyses of the value and utility provided by fractionated spacecraft or on design studies of specific systems. This thesis addresses two key questions: Firstly, is fractionation a mission enabling technology? Secondly, can design rules and guidelines be developed for fractionated satellites to allow continuity of measurements to be maintained and launched mass minimised? The first question is addressed by studying a payload and system for measuring coastal salinity from space. This provides an initial opportunity to assess the fractionated concept from a systems engineering point of view. The second question is addressed by undertaking a more general analysis of fractionated architectures, building on the knowledge gained in development of the fractionated coastal salinity measurement system. A computer model was developed to simulate the lifetime of different fractionated architectures when subjected to subsystem failures. A local search algorithm was used to find fractionated architectures which gave the best compromise between mass launched versus the operational time over a 50 year lifetime. The results from the application of this model showed that architectures that are highly fractionated, containing several homogeneous satellites, best achieved this compromise. These findings provide a contrast to the heterogeneously fractionated architectures proposed by DARPA. When a fractionated architecture is first implemented, the technology required to fractionate all the spacecraft subsystems may not be available. Consequently, the key to the implementation of these first fractionated architectures will be to ensure that there is redundancy in the fractionated subsystems spread across the architecture.

The present trend in spacecraft is to have two separate thrusters systems performing different tasks, a main electric propulsion system operating on xenon and a chemical system, usually bipropellants or cold gas. The development of a low power electric propulsion system operating on xenon to replace the chemical thrusters on board spacecrafts would be beneficial. It would be bring significant advantages in terms of mass saving from the sharing of the tanks, pipes and flow control unit, also with improvements in the specific impulse. In recent years experiments have demonstrated the possibility of using hollow cathodes as standalone thrusters, with indirect thrust measurement performed at the University of Southampton. Nevertheless indirect thrust measurements bring large uncertainties on the real value of the thrust. For the first time, direct thrust measurements were carried out with two different thrust balances on two different hollow cathode thrusters, derived from the T5 and T6 hollow cathodes, with unique design modification in the orifice and anode geometry. These measurements provide a unique insight into the real performance range of hollow cathode thrusters. Significant improvements in thrust, specific impulse and thrust efficiency have been achieved thanks to the optimized design of the T6 hollow cathode. The design of the thruster was modified using a one dimensional theoretical model developed within this research. With the help of the theoretical model the optimisation of the hollow cathode thruster design was carried out and a better understanding of the physical mechanisms which contribute to the generation of the thrust could be achieved, with the conclusion of electrothermal and electromagnetic phenomena being the main contributors. The main conclusions of the research and recommendations for related future works are also presented.

[by] Gordon K. Mandell, George J. Caporaso [and] William P. Bengen. / Includes bibliographies.

Deleterious effects such as surface heating and turning moments can arise as a consequence of the impingement of thruster plumes with spacecraft surfaces. Such thrusters are normally fired for attitude control. The prediction of such effects must be undertaken at the design stage of the vehicle. In this study, the modelling of such plumes was undertaken. The following prediction techniques were implemented into computer programmes: (a) the Simons model, (b) the Method of Characteristics (MOC), and (c) the Direct Simulation Monte Carlo method (DSMC). The first two methods are derived from continuum equations whilst the third adopts a discrete particle approach. Several DSMC schemes exist for treating the collisional behaviour of the gas, and it was unclear which would be best suited for the intended application. A thorough assessment of the implementation and performance of several such schemes was therefore completed. Having determined the most suitable DSMC scheme, the three modelling techniques were then applied to the isentropic core expansion of a small hydrazine thruster plume. It was found that significant errors occur in the determination of impingement quantities through application of the continuum methods in the flow regime lying between the continuum and free molecular limits. The DSMC technique was also used to calculate the nozzle lip and backflow expansion regions of the same hydrazine thruster. A significant degree of backflow was found with flow angles of up to 140º. The sensitivity of the calculations to the conditions initially assumed were assessed and found to be important.