Many control problems are based on control objectives easily quantified and consequently realisable by standard control synthesis methods. When an unmanned aircraft navigates, it moves inside a complex environment due to interactions with its surrounding and time varying environmental conditions. Several causes of perturbations have been identified as for example gusts and corrupted information of position. The characteristics of possible missions carried out by the un manned aircraft leads to the desire to construct navigation control systems which when operated in perturbed environments combine the advantages of smooth control with accurate navigation. Rule based, and adaptive controllers have favourable properties for such systems. This thesis investigates the use of a rule based navigation controller for a particular unmanned aircraft, the XRAEl aircraft. To achieve this objective several different types of fuzzy logic controllers are analysed as for example conventional and direct and indirect adaptive fuzzy controllers. They are designed by employing simple control engineering knowledge and subsequently validated using a stability method. For this purpose diverse stability methods are described and a new technique presented, the fuzzy root locus method, which is also based on the introduction of a new concept for fuzzy logic controllers, the fuzzy cell. The realisation of this work has been achieved by a series of simulation tests employing different processes and a simulation model of the XRAEl aircraft. The conclusions drawn from the results of the experiments consider in general that a rule based controller can improve the quality of navigation when compared to conventional controllers.

The development of a microprocessor based digital flight control system for a particular R.P.V. is described. The tasks required of this system are defined, and thereafter, the hardware circuits and the software structure necessary to implement a prototype are presented. The autopilot control laws are inferred from z-plane root loci, and then confirmed using digital simulations of the de-coupled roll and pitch attitude loops. The problems of the finite wordlength implementation of the control laws are discussed, and then both hybrid simulation and actual flight results are used to prove the performance of the prototype. To exploit the adaptive capabilities of a software based system, a sliding mode variable structure control law is developed for the roll attitude loop. Digital simulations are used to show that significant improvements in sensitivity reduction can be achieved under some conditions. These improvements are lost if a realistic servo-actuator model is employed. Another objective, namely the reduction of the disturbance error induced by trim imbalance, is maintained provided a reduced order switching function is used.

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Thesis (M.S.)--Brigham Young University. Dept. of Mechanical Engineering, 2005. / Includes bibliographical references (p. 55-58).

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A study has been conducted of reaction controls for VSTOL aircraft using thrust augmenting ejector techniques. Rapid mixing nozzles have been developed for high pressure ejectors. Mass flow increases for sonic nozzles of up to 50\ at x/D=8 were recorded, compared with plain circular nozzles. Their use was found to improve the thrust performance of a simple ejector by 9\, and larger increases are believed possible. Results from an ejector performance prediction model were successfully compared with experimental data. The use of rapid mixing nozzles in a practical ejector design has been assessed. It is predicted that a maximum thrust increment of 20\ ·could be achieved, compared with a simple fully expanded jet flow.

Integrated Modular Avionics standardises hardware and software platforms of Line Replaceable Modules (LRMs) and other system components in order to reduce the overall cost of system development. operation and maintenance. Several identical processing units within a cabinet. and fast communication media in the form of a backplane bus introduces further possibility of reconfiguring the system in terms of changing the applications performed by particular core LRMs. In this thesis a study into Reconfigurable Integrated Modular Avionics is presented. The main objectives of the project were to investigate the benefits, and feasibility of, employing autonomous dynamic in-tlight reconfiguration of the system as a means for providing fault-tolerance. In this approach, allowing processing modules to change their function permits the system to share the redundant modules as well as sacrificing less important avionics functions to sustain the more critical applications. Various architecture examples are reviewed in order to establish a system design that would support reconfiguration at a minimal cost. Two modified ARINC 651 architecture examples are proposed for implementation of dynamic in-flight reconfiguration. The benefits of reconfiguration are identified with the use of Markov state space analysis, and are found to be substantial with respect to the reduced number of redundant processing modules required to implement the system functions within the safety requirements. Suitable reconfiguration schemes are identified, and the most promising one is formally specified with the use of the Vienna Development Method. The safety properties of the scheme are shown based on the specification. In order to study the feasibility of autonomous dynamic reconfiguration, the scheme is implemented into two distinct systems, and the results of the practical observation of the system behaviour are presented and discussed. As the project was sponsored by the UK Civil Aviation Authority, a number of certification issues related to reconfigurable avionics systems are identified and discussed based on the practical implementation and previous theoretical analysis. It is concluded that dynamic in-flight reconfiguration of avionics systems can lead to substantial savings in terms of the reduced number of required core LRMs, and greater fault-tolerance than traditional non-reconfigurable systems

Any redundancy scheme in aircraft control systems is usually considered separately from the control algorithms involved. All feedback control systems are usually designed under the assumption that their sensors will not fail. When the integrity requirements demand it, then a redundancy scheme must be designed to provide any required measurements with only extremely short interruptions to normal service being caused by failures of individual sensors.