This thesis reviews the developments in commercial aviation resulting from More Electric Aircraft initiatives. The present level of electromechanical actuation is considered with discussion of the factors affecting more widespread use. Two rather different electromechanical actuators are presented for commercial aircraft; DEAWS electrical flap actuation and ELGEAR nose wheel steering. Both projects are industrially driven with specifications based on existing medium-sized commercial aircraft. Methods comparing fault tolerant electric drive topologies for electrical actuators are presented, showing two different categories of electric drive and comparing each category in a variety of operating conditions to assess size and component count. The safety-driven design process for electromechanical actuators is discussed with reliability calculations presented for both proposed actuators, showing where fault tolerant design is required to meet safety requirements. The selection of an optimum fault tolerant electric drive for each actuator is discussed and fault tolerant control schemes are presented. The development of the electric flap and nose wheel steering systems is described, with the focus on the work performed by the author, primarily on the power electronic converters and control software. A comprehensive range of laboratory and industrial results are given for both actuators, showing demonstrations of fault tolerance at power converter and actuator levels. Following testing, further analysis is given on various issues arising prior and during testing of both converters, with design considerations for future electromechanical actuators. From design testing and analysis, the two projects can be compared to attempt to determine the optimal electromechanical actuator topology and to consider the challenges in evolving the two actuators to aerospace products.

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The optimisation of aerospace structures is a very complex problem, due to the hundreds of design variables a multidisciplinary optimisation may contain, so that multilevel optimisation is required. This thesis presents the recent developments to the multilevel optimisation software VICONOPT MLO, which is a multilevel optimisation interface between the well established analysis and design software packages VICONOPT and MSC/NASTRAN. The software developed is called VICONOPT MLOP (Multilevel Optimisation with Postbuckling), and allows for postbuckling behaviour, using analysis based on the Wittrick-Williams algorithm. The objective of this research is to enable a more detailed insight into the multilevel optimisation and postbuckling behaviour of a complex structure. In VICONOPT MLOP optimisation problems, individual panels of the structural model are allowed to buckle before the design load is reached. These panels continue to carry load with differing levels of reduced stiffness. VICONOPT MLOP creates new MSC/NASTRAN data files based on this reduced stiffness data and iterates through analysis cycles to converge on an appropriate load re-distribution. Once load convergence has been obtained with an appropriate criterion, the converged load distribution is used as a starting point in the optimisation of the constituent panels, i.e. a new design cycle is started, in which the updated ply thicknesses for each panel are calculated by VICONOPT and returned to MSC/NASTRAN through VICONOPT MLOP. Further finite element analysis of the whole structure is then carried out to determine the new stress distributions in each panel. The whole process is repeated until a mass convergence criterion is met. A detailed overview of the functionality of VICONOPT MLOP is presented in the thesis. A case study is conducted into the multilevel optimisation of a composite aircraft wing, to demonstrate the capabilities of VICONOPT MLOP and identify areas for future studies. The results of the case study show substantial mass savings, proving the software's capabilities when dealing with such problems. The time taken for this multilevel optimisation also proves the efficiency of the software.

This thesis presents a time-conservative finite-volume method based on a modern flow simulation technique developed by the author. Its applicability to technically relevant aeroacoustic applications is demonstrated. The time-conservative finite-volume method has unique features and advantages in comparison to traditional methods. The main objectives of this study are to develop an advanced, high-resolution, low dissipation second-order scheme and to simulate the near acoustic field with similar accuracy as higher-order (e.g., 4th-order, 6th-order, etc.) numerical schemes. Other aims are to use a large-eddy simulation (LES) technique to directly predict the near-field aerodynamic noise and to simulate the turbulent flow field with high-fidelity. A three-dimensional parallel LES solver is developed in order to investigate the near acoustic field. Several cases with wide ranges of flow regimes have been computed to validate and verify the accuracy of the method as well as to demonstrate its effectiveness. The time-conservative finite-volume method is efficient and yields high-resolution results with low dissipation similar to higher-order conventional schemes. The time-conservative finite-volume approach offers an accurate way to compute the most relevant frequencies and acoustic modes for aeroacoustic calculations. Its accuracy was checked by solving demonstrative test cases including the prediction of narrowband and broadband cavity acoustics as well as the screech tones and the broadband shock-associated noise of a planar supersonic jet. The second-order time-conservative finite-volume method can solve practically relevant aeroacoustic problems with high-fidelity which is an exception to the conventional second-order schemes commonly regarded as inadequate for computational aeroacoustic (CAA) applications.

This thesis details the aerodynamic testing of a dielectric barrier discharge (DBD) plasma actuator operating over a separation step created at the trailing edge of a modified NACA 0012 aerofoil. The work focuses specifically on the use of pulsed or interrupted plasma actuation as opposed to continuously driven actuation, to increase the change in the lift produced by activating the system. The behaviour of the actuation system is characterised in a lamina flow regime at a Reynolds number of 1.33 x 105 using force balance measurements. At zero incidence the actuator produced a peak change in CL of approximately 0.015. However, this result is sensitive to changes in the interruption frequency of the plasma, by changing the plasma drive waveform the system was able to produce both positive and negative changes in lift. A relationship was identified between the change in CL produced and the ratio of the plasma interruption frequency to the natural vortex shedding frequency. This effect was investigated using both time averaged particle image velocimetry (PIV) and instantaneous phase locked PIV images captured in sequence throughout the plasma interruption cycle. The phase locked images showed how variation in the pulsation frequency was able to produce bi-directional actuation by either constructively or destructively interfering with the vortex formation from the back of the separation step. This interference in turn altered the level of separation which was occurring, altering the degree of upwash in the wake and therefore the lift generated by the aerofoil. PIV images were also gathered for device operation at a Reynolds number of 2.3 x 104; this produced a much higher ratio of DBD jet energy to that of the freestream. These conditions showed modified actuator behaviour due to the increased authority over the flow. However, the data still showed a strong interdependence on the reinforcement or destruction of the vortex street by the actuator interruption. Furthermore, work was undertaken to develop an actuator topology based on thin metallised films along with a dielectric which was hardened against the chemical and electrical stresses present in a functioning DBD device. The failure mechanisms of metallised film actuators were investigated, and actuators with lifetimes exceeding 8 hours were demonstrated. A manufacture method for a silicon polymer (PDMS) – Kapton® laminate is detailed; this is shown to be highly resistant to both electrical breakdown and chemical attack by the oxygen plasma.

The thesis documents the investigation of the application of the adjoint method to turbomachinery blading design optimization, with emphasis on blading aerodynamic design optimization in a multi-bladerow environment and concurrent blading aerodynamic and aeromechanic design optimization for a single bladerow. Based on the nonlinear flow equations, a steady adjoint system has been developed using the continuous adjoint approach. The capability of the conventional adjoint system has been augmented by the introduction of an adjoint mixing-plane treatment. This treatment is a counterpart of the flow mixing-plane treatment, enabling the steady adjoint equations to be solved in multi-bladerow computational domains. This allows turbomachinery blades to be optimised to enhance their aerodynamic performance in a multi-bladerow environment with matching between adjacent bladerows dealt with in a timely manner. The Nonlinear Harmonic Phase Solution method, a neat frequency domain method catered specifically for turbomachinery aeromechanics prediction, has been chosen to integrate with the adjoint method to calculate objective function sensitivities efficiently for concurrent aeromechanic and aerodynamic design optimization for single row turbomachinery blades. The Nonlinear Harmonic Phase Solution method, unlike the time-linearized methods, solves the unsteady flow equations at two or three carefully selected phases of a period of unsteadiness. This approach not only can conveniently turn a steady flow solver to one solving the unsteady flow equations efficiently, but also provides a good basis on which the corresponding adjoint system can be formulated and solved in a similar manner by extending a steady adjoint system. In order to resolve the issue of having a good blading performance over a whole operating range at a given operation speed, a multi-operating-point design optimization is implemented by formulating an objective function of a weighted sum of performance at more than one operating point

The thesis concerns the theoretical development and implemantation of sliding mode schemes for fault tolerant control. The theoretical ideas developed in the thesis have been applied to aerospace systems. In particular, actuator and sensor fault tolerant control schemes have been developed for a high fidelity full nonlinear model of a Boeing 747 aircraft which is a widely researched testbed in the open literature. A key development in this thesis considers sliding mode control allocation schemes for fault tolerant control based on integral action and a model reference framework. Unlike many control allocation schemes in the literature, one of the main contributions of this thesis is the use of actuator effectiveness levels to redistribute the control signals to the remaining healthy actuators when faults/failures occur. A rigorous stability analysis and design procedure is developed from a theoretical perspective for this scheme. A fixed control allocation structure is also rigorously analyzed in the situation when information on actuator effectiveness level is not available. The proposed scheme shows that faults and even certain total actuator failures can be handled directly without reconfiguring the controller. A design of an adaptive gain for the nonlinear component of the sliding mode controller for handling faults is also described. The later chapters of the thesis present the results obtained from real time hardware implementations of the controllers on the 6-DOF SIMONA flight simulator at Delft University of Technology as part of the GARTEUR AG16 programme. The schemes have been evaluated by experienced pilots and the results have shown good performance in both nominal and failure scenarios. A reconstruction of the Bijlmermeer ELAL 1862 scenario on SIMONA using one of the controllers shows that a safe flight and landing is possible with significant reduction in pilot workload when compared with the classical controller.

This thesis describes the application of H∞ design techniques to the control of high performance aero-engines. The design study presented is practical and realistic, the work being motivated by problems that arise naturally in real engineering situations. The aero-engine is multivariable and highly nonlinear: the dynamics vary considerably with the thrust being produced, and with the altitude and forward speed of the aircraft. Moreover, there are operational constraints that must never be violated for reasons of safety: certain engine variables should always be limited to safe vales. Furthermore, not all the engine parameters to be controlled are directly measurable; instead a number of related measurements are available. A methodology is presented to choose from the available measurements, those that are preferable for feedback control. Different techniques of model reduction using balanced realizations are considered. Two illustrative examples are presented, and the methods compared in detail. Explicit state-space formulae for an H∞-based two degrees-of-freedom robust controller are derived in discrete time. The controller provides robust stability with respect to coprime factor perturbations, and a degree of robust performance in the sense of making the closed-loop system match an ideal reference model. Special attention is paid to the structure of the controller. It is shown that the controller consists of a plant observer, the reference model, and a generalized state feedback law associated with the plant and model states. Multi-mode control logic is developed to ensure that safety limits are never violated. Actual engine test results are presented for sea-level static conditions. All the different modes of operation are tested. Full flight envelope evaluation of the controller is carried out using a nonlinear engine simulation. The robust performance of the controller is demonstrated and comparisons made with existing engine control systems.

This thesis is concerned with the results of a joint academic and industrial study on the development of a detailed nonlinear dynamic model of a turbofan jet engine to be used for research into advanced control strategies for civil turbofan aircraft engines. The model is representative of a dual shaft engine with variable bleed, variable stator vanes, turbine cooling, heat transfer, and a duct and exhaust nozzle. A switched, gain-scheduled, feedback control system incorporating bumpless transfer and antiwindup functionality has been designed and implemented according to current industrial practice. This baseline implementation permits realistic transient operation of the simulation and may act as a reference design for further control work. The simulation computes a non-iterative solution, by progressing calculations in the direction of the gas stream flow. Where possible the underlying physics are used and empirical approximations are avoided so that the model requires minimum data. This approach also makes a future inclusion of component failure easier to implement. The simulation is modular in nature so that engine or control modules can be easily replaced or modified if an improved design becomes available. The Simulink implementation of the control architecture has been redesigned to permit the addition or removal of control loops, also during the simulation’s operation, to allow testing of advanced control strategies. The entire controller can also be easily replaced. A detailed description of the modeling process, the various simulation issues that arise with a model of this complexity, and the results of the overall aero-engine system are presented. The design of the switched, gain-scheduled aero-engine controller with bumpless transfer and antiwindup which achieves dynamic performance that closely matches that of a real aero-engine is also discussed.