An experimental study into the phenomenon of aerodynamic noise generation in control valves has been performed. Several model representations of control valves have been studied, in addition to making measurements on several full size control valves. A high speed computer based data acquisition system, coupled with miniature high frequency surface mounted pressure transducers, has been used to obtain measurements of both the mean and fluctuating wall static pressures at several locations throughout the length of the flow apparatus. For known flow conditions, it has been possible to determine the relative performance of each particular model or commercial valve design by examining the various recorded levels of the wall static pressure fluctuations. A series of flow visualisation experiments, in which both wall static pressures and photographs of the flow through a rectangular duct have been obtained simultaneously, has provided a valuable insight into the mechanisms that determine the performance of multi-plate models. These findings have led to the design and commissioning of a new working section constructed from lengths of circular pipe. This apparatus has been used to investigate the performance of models constructed from two circular multi-holed test plates placed normal to the flow, with the downstream test plate being significantly more porous. Relationships between parameters such as test plate hole diameter, test plate separation, hole pitch to diameter ratio and pressure ratio have been examined extensively for many combinations of model configuration. Subsequently, the vast amounts of experimental data produced by this systematic testing have been reduced to identify the clear links between the various parameters. This, in turn, has enabled the author to determine optimum values for the non-dimensional parameters which govern the design of such multi-plate systems. As a consequence of this study, the author has proposed an alternative approach to the design of 'Low Noise' control valves which offers the prospect of improved overall performance. Thus, it is envisaged that the findings from this investigation will have implications for the future design of 'Low Noise' control valves.

This thesis contributes towards the discipline of whole aircraft simula- tion; modelling ight dynamics and airframe systems simultaneously. The objective is to produce estimates of the dynamic power consumption char- acteristics of the primary ight control actuation system when executing manoeuvres. Three technologies are studied; the classic hydraulic actuators and the electromechanical and electro-hydrostatic types that are commonly associated with the more electric aircraft. Models are produced which represent the ight dynamics of an aircraft; these are then combined with low frequency dynamic functional models of the three actuator technologies and ight controllers. The result is a model, capable of faster than real time simulation, which produces estimates of ac- tuator power consumption as the aircraft follows prede ned trajectories. The model is used to quantify the energy consumption as a result of di erent manoeuvre rates when executing banked turns. The result from an actuation system point of view alone is that the lower the turn rate, the lower the overall energy used. The tradeo is that the turn radius becomes larger. The use of the model can be extended to assist with additional design challenges such as actuator design and speci cation. Using methods to size actuators based on stall force and no load speed properties leads to oversizing of the control system. Performing dynamic analyses is usually a combined task of laboratory based actuator test rigs stimulated by input data gathered during ight tests. The model in this work provides a method of generating data for preliminary design; therefore reducing the amount of ight testing required in a design and certi cation programme. The major results discovered using the tools developed in this thesis are that a hydraulically powered aileron uses 4.23% more energy to achieve a turn at a heading rate of 0.03 rad/s compared to a 0.005 rad/s manoeuvre in the same conditions. The electromechanical actuator (EMA) uses 1.67% more and the electrohydrostatic actuator (EHA) uses 1.54% more to achieve the same turns. It implies reduced turn rate turns would have the largest bene t for reducing energy consumption in current hydraulically powered actuation systems, compared to electrical actuators.

An approach is proposed in this work to support the preliminary design of High-Lift aircraft configurations through the use of Multi-Objective optimisation tech¬niques. For this purpose a framework is developed which collates a Free-Form De¬formation parametrisation technique, a number of Computational Fluid Dynamics suites of different fidelity levels, a rapid aero-structure coupling procedure and two multi-objective optimisation techniques, namely Multi-Objective Tabu Search and Non-dominated Sorting Genetic Algorithm-II. The proposed optimisation framework is used for the execution of several design studies. Firstly, the deployment settings and elements' shape of the 2D multi-element GARTEUR A310 test case are optimised for take-off conditions. Consider¬able performance improvements are achieved using both the optimisation algorithms, though the sensitivity of the optimum designs to changes in operating conditions is highlighted. Therefore, a new optimisation set-up is proposed which successfully identifies operational robust designs. Secondly, the framework is extended to the optimisation of 3D geometries, using a Quasi-three-dimensional approach for the evaluation of the aerodynamic performance. The application to the deployment settings optimisation of the (DLF F11) KH3Y configuration illustrates that the method can be applied to more complicated real-world design cases. In particular, the deployment settings of slat and flaps (inboard and outboard segments) are suc¬cessfully optimised for landing conditions. Finally, a rapid aero-structure coupling procedure is implemented, in order to perform static aero-elastic analysis within the optimisation process. The KH3Y optimisation study is repeated including, this time, the effects of structural deformations. Different optima deployment settings are identified compared to the rigid case, illustrating that, despite being of reduced magnitude, wing deformations influence the optimum high-lift system settings. Furthermore, an industrial development and application of multi-objective opti-misation techniques is also presented. In the proposed approach, a reduced order model based on Proper Orthogonal Decomposition methods is used in an offline-online optimisation strategy. The results of the optimisation process for the RAE2822 single-element aerofoil and for the GARTEUR A310 multi-element aerofoil illustrate the potential of the method, as well as its limitations. The technical analysis is com-pleted with a description of the Agile project management approach used to run the project. Finally, future work directions have been identified and recommended.

Obtaining reliable CFD predictions of the bronchial flow that reflects the actual flow within a living lung requires the development of a deforming airways model, and the imposition of physiological subject-specific boundary conditions. This thesis addresses these two issues by the development of dynamic CFD models of the bronchial airways using a dynamic CT data set covering the breathing cycle of a laboratory animal. A deformation algorithm is proposed that matches the CFD mesh of the subsequent airway geometries generated from the dynamic CT data set. In addition, a novel nonlinear dynamic airway model generated from a pair of CT images is introduced. The proposed non-linear deforming model is capable of successfully capturing the non-linear motion characteristics of the bronchial airways based on the clinical measurements of the lung volume change. Furthermore, a technique to drive physiological subject specific boundary conditions for the terminal surfaces of the CFD models of the bronchial airways is introduced. The proposed technique depends on approximating the lung volume associated to each terminal surface over several time points over the breathing cycle based on the mechanical coupling between the bronchial airways and the vascular tree. The computed dynamic subject-specific boundary conditions were imposed on the terminal surfaces of the deforming airway model and the effect of wall motion on the flow features during tidal breathing is investigated for the first time. The outcome of this thesis is expected to improve the fidelity of the CFD predictions of the bronchial flow compared to the actual flow within a living lung. In addition, the availability of a new non-linear dynamic model of the bronchial airways that requires one pair of CT images as input, which complies with the radiation dosage restrictions for humans will facilitate the development of well-resolved CFD models of the human bronchial airways.

The subject of flapping wing Micro Air Vehicles '(MAVs) has been an area of increasing interest in the fields of aerodynamics and vehicle dynamics and control. The low Reynolds number region in which they operate, coupled with the periodic nature of the inputs to the control surface and the flexibility of the primary force-generating surfaces, has led to body models that are far from conventional, well established models used in the aerospace industry. It is therefore difficult to evaluate stability and sensitivity to design and operational parameters. The objective of the work in this thesis is to investigate the inherent nonlinear behaviour in flapping wing longitudinal flight to gain a deeper understanding of the stability attributes of the flight regime. To this end, a rigid flapping wing model was developed from first principles and analysed using continuation methods to determine the stability deviation with varied design and operating parameters. Additional continuation runs provided solutions trimmed to hover and forward flight, on which further stability and performance studies were carried out. The research revealed the existence of multiple steady-state solution branches within flapping wing flight and provided insight into the variations in behaviour of the nonlinear periodic system as input parameters vary. In particular, the III flapping frequency, the wing lead-lag motion, wing position and vehicle mass were shown to have the largest influence on stability. Further study revealed branches of stable and unstable limit cycles trimmed to hover and forward flight; analysis of which demonstrated that the areas of poor stability characteristics I correspond to low power requirements. Additionally, it was shown that previously stable areas in hover flight underwent per.io d doubling bifurcations with increased forward speed, leading to unstable and undesirable solutions. IV

Recently, the use of small-scale rotorcraft unmanned aerial vehicles (UAVs) for surveillance and monitoring tasks is becoming attractive. Their usage can be extended for monitoring of oil and gas pipelines. Amongst the various configurations of small-scale rotorcraft UAVs, the use of a quadrotor gained more prominence, particularly in the academic research community. A quadrotor is a small responsive four-rotor vehicle controlled by the rotational speed of its rotors. It is compact in design with the ability to carry a high payload. Achieving a successful outdoor surveillance/monitoring task with a quadrotor requires a robust and disturbance-rejection flight controller. However, to design a robust disturbancerejection flight control system for a quadrotor, the understanding of its aerodynamic permanence is essential. The fact that thrust is the key variable which influences the overall performance of rotary-wing air-vehicles, whilst wind velocity variation affects thrust utilization, their overall impact on a quadrotor manifests a rise or a fall in the rotors thrust which results in a non-uniform rotors’ thrust, and subsequent drop in altitude. This thesis addresses the effect of wind velocity variation in the design of a quadrotor autonomous flight control system. The following represents the main achievements recorded in this study: extensive review of the relevant literature related to UAVs with emphasis on rotorcraft, a quadrotor system was designed and developed in-house, followed by its aerodynamic analysis using the computational fluid dynamics (CFD) modeling technique, where ANSYS FLUENT and a virtual blade model (VBM) were employed to analyse the wind velocity variation effects on the quadrotor. A frequency-domain system identification technique was used to extract the quadrotor parameterized model using the comprehensive identification from frequency responses (CIFER) software package. Matlab/Simulink tools and Arduino-Simulink blockset were utilized in the design of the quadrotor’s altitude and attitude regulating controllers, based on the classical PID and model predictive control (MPC) schemes. Two algorithms were also developed, one for the vehicle’s navigation control and the other a graphical user interface (GUI) which facilitates semi-autonomous/autonomous control of the quadrotor. A series of actual flight tests were conducted to evaluate the performance and effectiveness of the quadrotor system. The tests involved careful selection of specific missions, based on which corresponding flight trajectories for the flight scheduling layer in the flight control structure were defined and the actual flight tests conducted. The flight scheduling involves ii the generation of appropriate flight trajectories for certain defined manoeuvres to evaluate the robustness and handling quality of the overall flight control system within the vehicle’s operational envelope. The defined manoeuvres include ascend-hover-descend, forward flight, hover turn and circular-zigzag flight. Each flight manoeuvre was defined in terms of clear objectives, full description and performance requirement, which was categorized into two qualitative levels, namely desired level (satisfactory) and adequate level (barely acceptable). Prior to conducting the extensive flight tests, the responses of the vehicle to specific directional control commands was examined using a frequency sweep excitation signal. The results of these tests were considered more than acceptable. This study lays a foundation for related researches, particularly in the development of small-scale rotorcraft flight control systems from an aerodynamic point of view.

This PhD study developed a composite structure with extreme orthotropic stiffness properties that offers the possibility for use as a skin panel in a morphing trailing edge control surface. The work has shown through an extensive literature review a lack of maturity in existing morphing skin concepts and thus the need for further research and development in this field. The focus of this investigation was on arranging conventional composite materials in a hierarchical structure that allows the combination of inherently different properties such as compliance and stiffness. This resulted in a fibre reinforced composite corrugated sandwich structure that is stiff parallel to the corrugation direction and relatively compliant, in tension and flexure, transverse to the corrugation direction. Experimental, analytical and numerical structural analysis was carried out to define the envelope within which corrugated structures can be designed to meet the requirements of a morphing skin. Further experimental and numerical aerodynamic investigations showed how to best implement these corrugated morphing skins in order to minimise the aerodynamic penalties. The combination of these results led to the design, manufacturing and testing of an aerofoil section with a morphing trailing edge control surface that incorporates a corrugated morphing skin. Low speed wind tunnel tests proved the concept but also highlighted limitations and raised suggestions for future work

Active control of flow separation over multi-element high-lift systems of aircraft wings could result in the decrease of drag force, the increase of lift, and the reduction of system complexity. This, in return, lowers fuel consumption and increases profitability, making the goals of both the Kyoto Protocol and the European Vision for 2020 more achievable. To maximise the potential of this technology, forms of actuation with high efficiency, low power consumption, fast response, good reliability, and low cost, are required. The synthetic jet is a kind of jet that on one hand produces zero net mass flux across the orifice over an oscillation cycle, whereas on the other is capable of transferring momentum and vorticity to the external fluid. It is a promising form of actuation for flow separation control with demonstrated success in laboratory experiments. The aims of the present research are to achieve an improved understanding of the fluid mechanics of synthetic jet actuators and to obtain the desirable modelling capability, such that a methodology of designing synthetic jet actuators for flow separation control at full-scale flight conditions can be developed. The research is focused on synthetic jets issuing from normal circular orifices. 2D axisymmetric numerical simulations were conducted using a commercial CFD code, FLUENT. The computational results for laminar synthetic jets agreed well with the experimental data, and the RNG 1(-cand Standard 1(-OJ models produced the best match between the computational and experimental results for turbulent synthetic jets. The CFD study confirmed the capability of FLUENT in simulating the key features of synthetic jets and established confidence in using the simulation results from FLUENT to validate low-dimensional models when experimental data are insufficient. Three low-dimensional prediction models have been developed, which are capable of predicting the space-averaged jet velocity from which the other actuator performance parameters can be estimated. The Dynamic Incompressible (DI) model provides analytical expressions for calculating the performance parameters for a given actuator geometry and operating condition. The Static Compressible (SC)model yields important relations among the actuator geometry and operating frequency, which allows the peak jet velocity to be maximised for a given diaphragm displacement, in two different flow regimes, i.e. the Helmholtz resonance regime and the viscous flow regime. The Lumped Element (LE) model is the most accurate model among the three. It is able to predict the jet velocity, which is in good agreement with CFD simulations, for both micro-scale and macro-scale actuators. An improved understanding in the effects of the dimensionless actuator operating parameters on the vortex rollup of synthetic jets has been achieved. The Stokes number S, the dimensionless stroke length L, and the Reynolds number based on the stroke length ReL, are important parameters affecting the vortex rollup. A criterion for the occurrence of vortex rollup in terms of the Stokes number has been developed. Finally, on the basis of the enhanced modelling capability and the improved understanding of the fluid mechanics of synthetic jet actuators, the first methodology of designing synthetic jet actuators for flow control at full-scale flight conditions has been developed. This methodology has been illustrated in the design of actuators on the leading and trailing edge devices of a multi-element high-lift system of a typical commercial aircraft wing at take-off conditions.

This thesis reports on an experimental investigation into the trailing edge pressure and losses of a symmetrical isolated aerodynamic body under transonic flow conditions. Tests were made in the 4" x 4" transonic wind tunnel at the University of Liverpool. The test programme included bleed through three different base configurations and the effect on base pressure and losses. Experimental measurements included base pressures, static pressure variation over the model surface and wake total pressure traverses. Schlieren photographs were taken and samples of these are presented to indicate the nature of the flow in the wake region.

The objectives of this PhD were to investigate the aerodynamic performance effects due to runback ice accretion with particular interest on the EASA 45 minute hold case in icing conditions. The sponsors, Airbus and Cranfield University collaboratively identified this aim as a result of the successful creation and capture of full-scale runback ice using the Cranfield Icing Tunnel. The hold phase represents typical icing conditions but with increasing demands on airports and subsequent knock-on effects to increased holding times has led to airliners paying more attention to this phase of flight. Typical icing conditions occur during the hold phrase of flight and place increasing demands on airports. A further challenge is that the EASA 45 minute hold case fails to take into account the large supercooled liquid droplets (SLD) when certifying the airplane. Published literature open to the public domain on flowfield interaction with realistic runback ice shapes, force coefficient losses and heat transfer interaction is limited. This is despite the fact that these parameters play a significant role in the examination of ice accretion. Inspection of the runback ice casting highlighted regions of two-dimensional features which were used for aerodynamic analysis. These high-fidelity two-dimensional runback ice shapes were utilised throughout this project. A single hex-core hybrid mesh from an ANSYS ICEMCFD script was designed and served the dual purpose of assisting the process of optimisation and validation. The numerical validation procedure analysed three separate studies with differing airfoils. The three studies examined were B737-700, B737- 200ADV and NACA 23012. The first two studies were in clean cruise configuration and the third simulated forward-facing quarter-round ridge ice. The experimental validation process investigated the drag associated for the three run back ice shapes. Tests were conducted in the atmospheric boundary layer wind tunnel. A multi-objective Tabu Search optimiser was coupled with ANSYS FLUENT solver for investigations on ice location, shape optimisation using freeform deformation, vertical tail plane shape optimisation and leading edge heattransfer effects. The primary finding in relation to ice location optimisation was an observed sensitivity to chord location. Further exploration of this data identified two district characteristics for this sensitivity. These characteristics are the variation of run back ice height relative to the size of boundary layer thickness. More specifically, if the boundary layer thickness is smaller than the runback ice height, trends identified from this project are adhered to. However, if the boundary layer thickness is larger than the run back ice height, no visible trends are observed. The ice location study found optimum chord locations closer to the leading edge of the airfoil and sensitivity to small chord movements. The B737 shape study was conducted as a preliminary optimisation test to highlight best practices for shape optimisation using free-form deformation. The results showed more stringent requirements for geometrical constraint handling when dealing with shape optimisation. Small changes to the NACA vertical plane improved both clean and iced performance. This was attributed to the leading edge design producing a lower velocity around the airfoil. This was a highly constrained optimisation problem where improved results required an exhaustive and competent search mechanism. The MOTS code was best placed to conduct this search and was therefore used throughout this study. Anti-Icing optimisation incorporated heat transfer providing a more complete runback ice optimisation study. This more complete design comprised a multipoint optimisation code with three objectives and one variable; to minimise leading edge temperature output, maximise lift and minimise drag with variable heat input. The findings corroborate the results seen in the ice location optimisation study. The ice location study would benefit from an increased range of variability to observe the development of the trends found. Solver prediction capabilities for ice accretion studies would benefit from experimental data on roughness parameters associated with icing. The two-dimensional airfoil simulations would benefit from further CFD runs using three-dimensional swept wings. The Pareto-optimal designs are ideal candidates to compare the flowfield inadequacies alluded to by literature. As an interim study before threedimensional simulations are conducted, an extension to the cruise configuration airfoil study would be to deploy high-lift profiles. It would provide insight into ice accretion scaling methods by running a second NACA tail plane optimisation run using runback ice shapes scaled based on boundary layer thickness. Findings would provide invaluable information on the interaction of the two scaling methods with the boundary layer and how changes to flowfield characteristics effect force coefficients. Finally, further aerodynamic investigations with wind tunnel testing of representative full-scale runback ice shapes on a sectioned full-scale swept wing would be beneficial. This would complete the years of commitment and innovation by Cranfield University, sponsors and various students on the issue of ice accretion, particularly runback ice.