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Among current techniques, used to assist the modelling of atmospheric processes, is an approach involving the balloon or aircraft launching of radiosondes, which travel along uncontrolled trajectories dependent on wind speed. Radiosondes are launched daily from numerous worldwide locations and the data collected is integral to numerical weather prediction. This thesis proposes an unmanned air system for atmospheric research, consisting of multiple, balloon-launched, autonomous gliders. The trajectories of the gliders are optimised for the uniform sampling of a volume of airspace and the efficient mapping of a particular physical or chemical measure. To accomplish this we have developed a series of algorithms for path planning, driven by the dual objectives of uncertainty and information gain. Algorithms for centralised, discrete path planning, a centralised, continuous planner and finally a decentralised, real-time, asynchronous planner are presented. The continuous heuristics search a look-up table of plausible manoeuvres generated by way of an offline flight dynamics model, ensuring that the optimised trajectories are flyable. Further to this, a greedy heuristic for path growth is introduced alongside a control for search coarseness, establishing a sliding control for the level of allowed global exploration, local exploitation and computational complexity. The algorithm is also integrated with a flight dynamics model, and communications and flight systems hardware, enabling software and hardware-in-the-loop simulations. The algorithm outperforms random search in two and three dimensions. We also assess the applicability of the unmanned air system in ‘real’ environments, accounting for the presence of complicated flow fields and boundaries. A case study based on the island South Georgia is presented and indicates good algorithm performance in strong, variable winds. We also examine the impact of co-operation within this multi-agent system of decentralised, unmanned gliders, investigating the threshold for communication range, which allows for optimal search whilst reducing both the cost of individual communication devices and the computational resources associated with the processing of data received by each aircraft. Reductions in communication radius are found to have a significant, negative impact upon the resulting efficiency of the system. To somewhat recover these losses, we utilise a sorting algorithm, determining information priority between any two aircraft in range. Furthermore, negotiation between aircraft is introduced, allowing aircraft to resolve any possible conflicts between selected paths, which helps to counteractany latency in the search heuristic.

A fuel cell is a device that converts energy in the fuel and reactant into electrical DC power. Fuel cell powered aircraft are generally characterised by a low power to weight ratio (W/kg). The propulsion system of an unmanned aircraft needs a large range of power and fast response to fulfil the requirements of different flight phases and to balance the variations in the load demand. A proton exchange membrane (PEM) fuel cell is considered as a potential power source for high altitude UAS (unmanned aircraft systems) operations. At altitudes in excess of 10 km, very low atmospheric temperatures and pressures, and unexpected variations in the load demand put severe stresses on the operation and performance of PEM fuel cells. A stable and robust controller and fuel supply system that can provide fast and sufficient flow of hydrogen and air/oxygen to the reaction of the fuel cell is one of the critical objectives. In this research, a simplified mathematical model of the PEM fuel cell stack system is developed and validated with the commercially available 1 kW PEM fuel cell stack (H-1000) developed by Horizon Fuel Cell Technologies. Matlab-Simulink is used to implement the necessary design and simulations under various operational conditions. The implications of high altitudes on the operation and performance of a PEM fuel cell stack are investigated, and a PID controller is adopted to efficiently optimise and provide a sufficient flow of hydrogen and air/oxygen to the stack, in particular maintaining the flow rates of the reactants was deemed most critical at high altitudes operation. Also, in order to store the required oxygen and hydrogen, the design of storage vessels is considered. This research presents a design of a PEM fuel cell power system for unmanned aircraft systems with an integrated approach that enables estimation of required power for high altitudes UAS operation which is then used to determine the size and weight of the combined power-plant of fuel cell stack with hydrogen and air/oxygen vessels and the propulsion system of the UAS. This approach takes into the consideration the power capacity of fuel cell stack and the flight endurance as two main factors in designing the size and weight of storage vessels, and hence determining the overall weight of the UAS, which is a key requirement in the preliminary aircraft design phase. One of the research outcomes shows a potential in extending the flying duration and altitude for more than five hours and a half, reaching up to 11 km altitude, for a UAS with an overall weight of 32 kg, including a payload capacity of 2 kg, based on a 1 kW PEM fuel cell propulsion system.

The more-electric aircraft (MEA) concept is widely viewed as the next evolutionary step towards enabling the industry goal of developing optimised, fuel efficient aircraft. MEA have an increased dependency on electrical energy for distribution to secondary systems and, in order to service this increased dependence, the electrical power systems (EPS) are more complex with increased voltage distribution levels, power conversion stages and safety critical components compared with their conventional counterparts. These complexities will only increase in future platforms as they further embrace the MEA concept - the migration to increasingly novel, critical and complex EPS will incur several development and integration challenges. This thesis considers the fundamental challenge of maintaining high reliability standards within future aircraft EPS through the development of accurate and discriminative real-time protection systems which will react during fault conditions. Specifically, the thesis researches novel methods that improve real-time aircraft EPS protection and health management systems by 1) accurately diagnosing degraded faults before their progression to critical failure and 2) diagnosing faults that are difficult to detect using only conventional protection methods – in particular, series arc faults are considered. Within future aircraft EPS, the volume of operational data is expected to significantly increase beyond that of the conventional systems; consequently, the thesis focuses on the use of data-driven, machine learning based methods, to enable these extended functionalities of the EPS protection and health management systems. The types of machine learning modelling techniques that were chosen are explained and justified. Conventional protection methods are described, including a discussion on the difficulties in using them to detect both degraded fault modes and arcing conditions. The necessity to detect these types of faults in an accurate and timely manner is also discussed. One of the main contributions of the thesis is the proposal of the EPSmart method that can autonomously diagnose and isolate a multitude of degraded faults within an aircraft representative EPS. These degraded faults include intermittent and incipient conditions, which, in comparison to overcurrent faults, often lack the energy to be detected by conventional means. Early, and accurate, detection of these conditions will improve overall system health management and reliability and ensure safe operation of the aircraft. Further contribution is the design of the IntelArc method that can detect series arc faults within direct current supplied systems. Accurate detection of series arc faults is extremely challenging as, despite their presence being a serious fire hazard, they result in a decrease of load current. Although methods do exist for diagnosis of series arcing, there remain challenges with regards to accurate detection across different system configurations and operating conditions. The thesis shows the potential for IntelArc to provide accurate detection across a variety of configurations and operating conditions. While the thesis only describes the initial development of these novel methods, the significant conclusions are that application testing has shown the potential for them to enhance real-time network protection, fault tolerance and health management of aircraft EPS through detection of degraded fault and arcing conditions.

In the conventional helicopter, the transmission and powerplant systems are the major production and operator direct operating cost drivers. Additionally, they are related to helicopter safety and reliability concerns and impose performance boundaries. As a consequence, they need to be addressed. Adapting the transmission and powerplant systems with the introduction of more electric technologies, which are reported to be more reliable and cost friendly, involves a gross weight penalty, which cannot be accepted from a performance viewpoint. The implementation of liquid hydrogen, with the objective to introduce a sustainable energy carrier and free cold source for the high temperature superconductive devices driving the tail rotor, appears unattractive, from either a weight or a exploitation standpoint. Biodiesel could be an alternative to Avgas driven configurations, but at the moment, it has questionable chemical characteristics and is therefore discarded. Conceptual alternatives to the conventional helicopter explored in an attempt to verify their ability to overcome the stated performance, safety and cost aspects, are subjected to the same problems, the Turbine Driven Rotor (TDR) helicopter configuration excepted. The TDR helicopter drives a coaxial rotor configuration by means of a rotor embedded Ljungstr¨om turbine, omitting the need for a mechanical transmission system. Three TDR helicopter thermodynamic cycles are proposed. The piston engine powered TDR cycle shows to be of interest for the low weight class helicopters. The turbofan powered TDR cycle is preferred in the mid and high weight categories, benefitting from its configurational simplicity. The more complex turboshaft powered TDR cycle requires a heat exchanger, is heavier and thus not recommended. With respect to the Ljungstrom turbine, the loss models of Soderberg and Ainley and Mathieson used to establish its geometry and performance characteristics are generally acceptable and appear coherent, while a deviation angle correction is developed to cope with the radial outflow configuration of the turbine. Similarly, loss models for the internal leakage and disk friction are proposed. However, these models could not be substantiated by means of experiments. A design methodology to implement the Ljungstrom turbine in the helicopter rotor head is presented and allows adjusting the thermodynamic cycle characteristics such as to maximise the performance gain with respect to the conventional helicopter. For nominal operating conditions, ISA SLS, a VLR-class TDR helicopter shows to bear a performance gain of 10% over a conventional helicopter when equipped with an Avgas engine and 14% when a Diesel engine is used. Hereby, the cycle pressure ratio remained low, i.e. approximately 1.25, allowing a turbine polytropic effciency of 87%. An identical study with a NH-90-class TDR helicopter proved to offer a performance potential of 50% at a cycle pressure ratio around 1.6 and a turbine polytropic effciency of 90%. In all cases, the gas temperature at the inlet of Ljungstrom turbine remained below the rotor bearing temperature limit of 400 K.

A postbuckled, carbon fibre reinforced plastic (CFRP) wing box has been designed, manufactured and tested for an aerobatic light aircraft, the Cranfield Al. Methods of analysis have been evaluated including: i) Non-linear finite element analysis for the prediction o-f panel postbuckling. ii) A simpler technique based on an effective width method. This forms the core of a design program, 'oPTIMIST'. It predicts buckling loads, postbuckled reduced stiffness and overall column failure of co-cured hat stiffened panels. It then optimises the con-Figuration of a box beam for minimum weight. iii) The use of the effective width method allied to a large scale, linear finite element analysis. The work includes the development of a new method o-F construction for composite box structures. The wing skin sti-Ffeners and rib flanges are co-cured together. Integral slotted Joint features are formed in each part. The structure is then adhesively bonded together. A full description of the manufacture o-F the wing box is included. The structure was also tested in a specially designed rig. It was tested to ultimate design loads in: i) Positive bending to 13.33. ii) Negative bending to -96. iii) Pure torsion resulting from full aileron load. iv) Torsion with 96 bending. The compression panels were seen to postbuckle and recover in each load case. Results are compared with theory, and with the original aluminium Al wing. The structure is 257. lighter than its aluminium counterpart. Finally, suggestions are made for possible areas of further research.

Today's high competitiveness in the airline industry urges for the development of even more efficient transport aircraft. In many cases lower operating costs are the key to survival. Although the introduction of emerging advanced technologies has shown improvements both in safety levels and performance, with the associated reductions in costs, the search for more economical aircraft must also take into consideration changes in current design practice. The study of novel configurations is a contribution to this view. In this research project, advantage was taken from multidisciplinary synergism to design and optimise conventional and three-surface configuration commercial aircraft, to satisfy the same mission and operational requirements. An integrated conceptual design synthesis approach was employed where typical aeronautical disciplines, as well their complex inter-relations, were taken intoaccount. All these considerations, together with both cruise and field performance, and static stability and control requirements, resulted in different baseline configurations of the two concepts, although sharing the same fuselage and the same technology standard, but with different Maximum Take-off Weights (MTOW), lifting surfaces, turbo-fan engine sizes, and economics. After coupling the design synthesis program to a gradient based numerical minimization routine, optimisation of these designs was performed for minimum Direct Operating Costs (DOC) and minimum MTOW, and their performance and economics were compared on an equal basis. Trade-off studies were conducted on all aircraft for 1000 through 3000 NM design mission ranges while keeping the same fuselage size, lifting surface planform shapes and same static longitudinal stability margin (inherently stable designs), as obtained for the respective datum designs (Range = 1250 NM). Thus, using the same comprehensive design tool, built on the same primary assumptions, and using the same analytical methods and principles which include many real life considerations, a systematic and conýistent study of both design concepts was conducted. The potential merits of a realistic three-surface transport design were clearly established, when comparison was made with an equivalent mission conventional twin turbo-fan airliner. Within the usual limitations of any initial conceptual design study, it appears that the concept of the three-lifting surface transport can effectively improve in terms of performance and direct operating costs, when compared to conventional aircraft designed for the same operational environment and mission profiles, and may show a promising future.

As all engineering works are a blend of theory and empiricism, aircraft design, by its nature, represents a mixture of aircraft designer's knowledge obtained from aeronautical engineering disciplines and its usage combined with his experience. This means not only the application but also the integration of all the fundamental knowledge of aerodynamics, structure, propulsion, stability and control, operational and economic aspects, etc., based upon the designer's judgements and experiences. Thus the tasks involved in designing an aircraft configuration, without exception, show complex characteristics, considering the fact that aircraft configuration design means the integration of components such as lifting surfaces ( wing ), fuselage, power-plant, control surfaces ( tail or canard ), and undercarriage. The discrepancies and mismatches among the aircraft components make the configuration design iterative, repetitive, and thus time - consuming. Such complexities of configuration design processes often require compromise, through trial and error, to resolve conflicts between the major design areas. Moreover, it takes tens of years to become a experienced design expert whose sound judgement, based upon experience and profound knowledge, influences greatly the aircraft configuration design. The differences in judgements depend upon the designers' imagination and experience, and they are the cause of variations in aircraft configurations. Therefore, the efforts were made to overcome those difficulties which hinder the aircraft designer from making the task of configuration design more efficient, and further to assist the aircraft designer in getting an easy and interactive preliminary aircraft configuration without always relying upon design experts. Hence the current research project is directed at the development of an expert system for aircraft design. This involves the use of Artificial Intelligence and its programming language called PROLOG ( PROgramming in LOGic ). The research started from a thorough analysis of the major component design areas and has constructed an EXPERT SYSTEM to find out the efficient Control Mechanism which can search intensively for the solutions to design problems for all types of aircraft; civil and military, subsonic and supersonic, conventional and unconventional, etc. In addition, users can have access to the explanations of important items such as a design process, terminology, equations, and results. The explanation facility is one of the most important functions of Expert Systems. Partly due to the limit of computer capacity and partly due to the magnitude of laborious program execution at this stage, the system implementation has focused on the high - subsonic, conventional and jet transport aircraft categories. The approach taken was to find an efficient and effective control mechanism (i.e. an Inference Engine ), which integrated the Parametric Study, Wing Design, Fuselage Design, Engine Design, Tail Design, Undercarriage Design, Weight Analysis and Cost Analysis into a whole configuration system. The comparison between Expert System results and existing aircraft such as Boeing 747, Airbus 300 series, BAe 146 series, McDonnell Douglas MD series, etc., showed the permissible ranges of error to be within about 10 %. Such results enable the Expert System to claim that it can act as a useful design tool for the aircraft designer in the initial stage of aircraft configuration design. Finally, the author believes that the control mechanism devised for this Expert System can be used as a sound basis for extending the Expert System to include other types of aircraft and further to encompass spacecraft design, as the designer wishes.

In this thesis a methodology for designing wings for transport aircraft operating at high subsonic speed is investigated. Several methods are studied, including more accurate methods such as the computational methods. These are used as an addition to the semi-empirical methods. Several attempts have been made to build a computerised aircraft design in the past. Most of the conceptual aircraft design programs that are available are based on the semi-empirical method only. As faster computers become available, a method for designing a high subsonic aircraft wing is studied by including computational aerodynamic and computational structural analysis in the integration process. SPARV is used as the computational aerodynamic program and NASTRAN is used as the computational structural analysis program. The objectives of this thesis are to study a method of performing, the conceptual design of wings for transport aircraft operating at high subsonic speed and to demonstrate that aerodynamics analysis using, Computational Fluid Dynamics (CFD) and structures analysis using the Finite Element Method (FEM), can be coupled with the aircraft synthesis program in a seamless distributed computing environment. The achievement of these objectives is demonstrated by applying the methodology to specific wing design. This method has been validated and tested for transport aircraft operating at high subsonic speed, but application on military transports may also be valid. An example case study is presented in this thesis. Improvement of the method for future development is also considered in the thesis. These include the use of a more powerful computational aerodynamics package.

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