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A three dimensional implicit multiblock Navier-Stokes solver for hovering rotor vortical flow simulations has been developed. The governing equations used are cast in an attached blade rotating frame. Two formulations of the governing equations using the relative or absolute velocity as variables respectively are employed and investigated. The Osher's approximate Riemann solver is used for the convective fluxes evaluation. A modified MUSCL scheme is employed for improving the accuracy of the discretisation for the in viscid fluxes. A Block Incomplete Lower and Upper Decomposition (BILU) is adopted for solving the linear system resulted from the use of an implicit scheme. Special treatment for the terms, including extra flux terms and source terms, arising from the non-inertial reference system are implemented. A multiblock technique is used to obtain the exibility for quality grid generation. The suitability of different grid topologies for vortex wake capturing is demonstrated. Numerical tests show that significant improvement in computational efficiency is achieved by utilising the BILU implicit scheme in both fixed wing and hovering rotor calculations. Numerical simulations also demonstrate Navier-Stokes solutions give more accurate results than that from Euler solutions, especially in transonic tip speed cases. Computed results including surface pressure distributions and tip vortex trajectories are compared with the experimental data, which shows that the developed solver and the numerical scheme can simulate hovering rotor flows with good accuracy.

This study investigates a novel methodology for the preliminary design of aeroengines. This involves the modelling of the disciplines that affect the engine's requirements and constraints, their implementation in software format and their coupling into a single unit. Subsequently, this unit is interfaced with an optimiser software. The resulting multidisciplinary optimisation (MDO) tool allows the automation of the traditional, human-based preliminary design process. The investigation of the above-mentioned novel methodology is carried out through the development of a "pilot" MDO tool and its subsequent utilisation in three case studies, characterised by different optimisation scenarios. The selection of each case study is motivated by current research questions, such as aviation's contribution to climate change or the attractiveness of specific novel propulsion concepts. The outcome of the pilot MDO study is considered successful and has been well received by several academic and industrial aero-engine organisations. The choice of the disciplines and of their modelling fidelity allowed a realistic representation of the main disciplinary interactions and tradeoffs that characterise the important phase of preliminary design. The computational effort involved in the solution of the optimisation studies was found to be acceptable, and no major reprogramming was required when different optimisation scenarios were considered. The case studies were investigated with an ease and comprehensiveness that would not have been achievable through a human-based parametric analysis. The positive experience with the pilot MDO tool suggests that an automated methodology for the preliminary design of aero-engines is feasible, applicable and valuable. Its adoption can provide substantial advantages over the traditional human-based approach, such as a reduction in human effort, costs and risk. From this perspective, the pilot study constitutes a first step towards the development of a full-scale MDO tooL usable by aero-engine manufacturers. In the near future, issues like climate change could drive significant modifications in airframe and engine design. A preliminary design MDO tool is therefore timely, and has the potential of making a significant contribution.

This thesis describes the development of a generic design methodology for variable camber flap systems for transport aircraft, intended to be used for cruise and low-speed operations. The methodology was structured after several revisions were performed on conventional high-lift device design methodologies for existing transport aircraft. The definition and detail explanations are given at every phase of the methodology. A case study was performed in order to give an example of the implementation of the methodology where a transport aircraft called A TRA, a design study from previous PhD report, was taken as a model. Experimental work could not be performed, due to budget constraints, so the case study was only carried out using computer-based analyses. Software packages such as MSES-code (a Computational Fluid Dynamic software), CATIA and PATRANINASTRAN were used for this case study to analyse aerodynamic characteristics, layout as well as simulation and structure analyses respectively. The results obtained showed that it was practically feasible to deploy such a high-lift device to transport aircraft when the effect from aerodynamic loads gave minimum effect on structural deformation. The deflections of the flap as well as spoilers under critical loads were below the allowable limits, which had a minimal effect due to the additional lift force generated from the movable surfaces.

The hovering flight of a micro air vehicle was investigated. Various flight control concepts, inspired by observation of insect and bird flight, were investigated in some detail. The concepts make use of flap frequency or phase between the flap and pitch attitudes of the wing to control the force magnitude. Tilt of the stroke plane or fuselage was used to vector the force. A flight control system was designed for each concept investigated. The investigation has revealed that the preferred control concept is one in which force vectoring is achieved by the fuselage tilt through a shift in the centre of gravity location while the force magnitude is controlled by the phase between pitch and flap attitudes. This has the advantage of reducing the vehicle weight while at the same time relaxing the demand of extremely high frequency actuators. The equations of motion based on a multi-body representation of a flapping wing micro air vehicle were derived and these form the basis for the SIMULINK flight simulation program used to carry out the above investigation. The aerodynamic model used for the simulation was obtained from force measurements with a flapping mechanism that allows the model wing to oscillate with two degrees of freedom (flap and pitch). During these measurements, the phase angle between the flap and pitch angles of the model wing was varied to determine the effects on the force magnitude and direction.

The purpose of this research has been to develop a technique by which satellite constellations in different classes of orbit may be realistically compared. Previous work on constellation design has tended to focus on minimising the number of satellites required to provide coverage of the Earth. The variations in satellite vehicle design, which result from the use of different orbits, have, in general, been neglected in such analyses. The purpose of this research is to bridge this gap between constellation design and satellite design using a Figure of Merit. This Figure of Merit incorporates the coverage value provided by the satellite constellation, measured in terms of percentage coverage time, and the overall mass of the satellites which are required to provide this coverage. The coverage value is measured against a specific requirement, which is defined geographically, and which may be weighted by the user to reflect the relative importance of different regions. This allows arbitrary, asymmetric, real-world requirements to be adequately represented. This also marks something of a departure from previous work, in that the goal of much constellation design work has been to provide un-weighted coverage of the entire globe. Simplified mass models are developed for generic communications and surveillance satellites in a variety of orbits, and are then used to calculate the Figure of Merit for individual satellites. It is shown that the best solution depends crucially upon the geographical distribution of the requirement, and other user-defined parameters, such as the minimum elevation angle which can be tolerated. It is also shown that, for certain typical requirements, the Figure of Merit correctly identifies geostationary orbit (GEO) and low Earth orbit (LEO) as having particular advantages. iii The technique of characterising the requirement geographically may also be used as a means of optimising the orbital parameters of the candidate constellations, and a preliminary description of this procedure is also provided. The Figure of Merit Technique is then applied to representative communications satellite constellations in order to demonstrate its ability to differentiate between candidate options. The Figure of Merit technique is also used to investigate the possibility of using a surveillance satellite at very low altitudes.

Ice growth may rapidly degrade the aerodynamic performance of an aircraft. It can also severely damage structures such as communication towers or power lines. Subsequently, de-icing and anti-icing systems have been developed and a number of codes designed to predict ice shapes. When ice accretion starts, two different types of ice can appear, depending on the temperature and conditions. All of the incoming fluid may freeze almost instantaneously and turn into rime ice. Alternatively, a fraction of the incoming fluid may freeze and turn into glaze ice while the other part remains liquid and may flow over the ice. Previous work on ice accretion has mainly targeted the ice shape and neglected the owing water layer. The present study focuses on this. A set of governing equations is derived for both rime ice growth and coupled ice growth and water flow. When rime ice accretes, a mass balance is used to calculate the shape. In the presence of both ice and water, the ice growth is governed by an energy balance and the water flow by a mass balance. These equations are solved numerically for the water flow alone and the coupled ice growth and water flow for two- and three-dimensional at inclined planes. The behaviour of both ice and water is studied. The model is then extended to deal with arbitrary substrates and solutions are sought for industrially important applications such as ice accretion on power lines or aerofoils. This research work forms part of the ICECREMO project. ICECREMO is a three-dimensional ice accretion and water flow code developed collaboratively by DERA, British Aerospace, Rolls Royce, GKN Westlands Helicopters and Cranfield University under the auspices of the UK department of Trade and Industry.

Usability is a multidimensional concept, and although a number of techniques have been developed for assessing usability, none have been developed specifically for the civil flight deck or its components. The requirement for Human Factors tools to support civil flight deck certification have been well documented (e.g. FAA, 1996). This thesis reports on a research programme that resulted in the development of a series of scales to assess the usability of Flight Management System (FMS) Control and Display Units (CDUs). Since the goal of this research was to produce scales that could be used by certification authorities and FMS manufacturers, demonstrating that the scales had appropriate psychometric properties was critical. A rigorous and systematic approach to the construction of the usability scales was therefore adopted. The research was conducted in three studies. The first study involved the development of a 49-item questionnaire that was distributed to airline pilots. A series of analyses on the data collected demonstrated that a reliable and valid usability factor structure was produced. The second study focused on the production of usability scales and examined the impact of flight phase on usability ratings. Analyses demonstrated that the factor structure produced in the first study was reliable, and was valid for data collected for the cruise and descent and approach phases of flight. A series of internally consistent usability scales was therefore produced which demonstrated evidence of discriminant, face, content, and construct validity. The third study involved a content analysis on a large volume of participant comments derived during studies one and two, and examined the usability problems associated with current FMSs and suggested improvements for future FMSs. The content analysis produced evidence of construct validity for the usability scales developed, and permitted recommendations for FMS designers and certification authorities to be made.

Helicopters are highly dependent on their transmission systems, which provide the vital links from the engines to the rotor and ancillary systems. Components are highly loaded and must be manufactured to a high degree of accuracy; the lack of redundancy implies that this is a 'series-chain' system. Existing techniques for calculating expected life are based upon historical data from different gearbox and helicopter types, thus limiting the confidence of the results. Design techniques may be conservative in some areas, whilst neglecting to consider different load patterns, usage, maintenance and environmental factors. This work describes the development of probabilistic models that represent damage accumulated by fatigue, wear and corrosion of the key components with an Intermediate gearbox (IGB). The parameters of these models represent geometrical, load and material data at the design stage, and produce an output in terms of failure probability against operating hours. This allows the influential parameters to be identified before building a prototype helicopter gearbox. The results from these models are then used to predict the upper and lower bounds of system reliability. This enables the combination of diverse failure mechanisms to be viewed to determine the relevant significance of each failure mechanisms. The effectiveness of the gearbox monitoring systems has been incorporated in the computer model by considering the probability of detection (POD) of each failure mechanism. The work to develop models found that there is a large body of work available to describe damage accumulation due to fatigue, but far less in regard to wear and corrosion. Fatigue models are very sensitive to load and material variability, particularly tooth root bending fatigue, for which many loads are considered 'non-damaging'. Wear models are mostly affected by changes in material hardness, wear coefficient and slip amplitude; changes in load are less influential on the predicted time to failure. The results for galvanic corrosion are dominated by the corrosion rate and time to initiate. In the system reliability model, reducing gear load appears to be the simplest means to increase life; increases in material strength and reduction in material variability are less achievable without significant improvements in manufacture and/or material technology.

The aim of this research is to develop an optimisation tool to support the preliminary design of gas turbine combustors by providing a partial automation of the design process. This tool is to enable better design to be obtained faster, providing a reduction in the development costs and time to market of new engines. The first phase of this work involved the analysis of the combustor design process with the aim of identifying the critical tasks that are suitable for being automated and most importantly identifying the key parameters describing the performance of a combustor. During the second phase of this work an adequate design methodology for this problem was defined. This led to the development of a design optimisation Toolbox based on genetic algorithms, containing the tools required for it's proper integration into the combustor preliminary design environment. For the development of this Toolbox, extensive work was performed on genetic algorithms and derived techniques in order to provide the most efficient and robust optimisation method possible. The optimisation capability of the Toolbox was first validated and metered on analytical problems of known solution, where it demonstrated excellent optimisation performance especially for higher-dimensional problems. In a second step of the testing and validation process the combustor design capability of the Toolbox was demonstrated by applying it to diverse combustor design test cases. There the Toolbox demonstrated its capacity to achieve the required performance targets and to successfully optimise some key combustor parameters such as liner wall cooling flow and NOx emissions. In addition, the Toolbox demonstrated its ability to be applied to different types of engineering problems such as wing profile optimisation.

An aerodynamic shape optimisation capability based on a discrete adjoint solver for Navier-Stokes flows is developed and applied to a Blended Wing-Body future transport aircraft. The optimisation is gradient-based and employs either directly a Sequential Quadratic Programming optimiser or a variable-fidelity optimisation method that combines low- and high-fidelity models. The shape deformations are parameterised using a B´ezier-Bernstein formulation and the structured grid is automatically deformed to represent the design changes. The flow solver at the heart of this optimisation chain is a Reynolds averaged Navier- Stokes code for multiblock structured grids. It uses Osher’s approximate Riemann solver for accurate shock and boundary layer capturing, an implicit temporal discretisation and the algebraic turbulence model of Baldwin-Lomax. The discrete Navier-Stokes adjoint solver based on this CFD code shares the same implicit formulation but has to calculate accurately the flow Jacobian. This implies a linearisation of the Baldwin-Lomax model. The accuracy of the resulting adjoint solver is verified through comparison with finitedifference. The aerodynamic shape optimisation chain is applied to an aerofoil drag minimisation problem. This serves as a test case to try and reduce computing time by simplifying the fidelity of the model. The simplifications investigated include changing the convergence level of the adjoint solver, reducing the grid size and modifying the physical model of the adjoint solver independently or in the entire optimisation process. A feasible optimiser and the use of a penalty function are also tested. The variable-fidelity method proves to be the most efficient formulation so it is employed for the three-dimensional optimisations in addition to parallelisation of the flow and adjoint solvers with OpenMP. A three-dimensional Navier-Stokes optimisation of the ONERA M6 wing is presented. After describing the concept of Blended Wing-Body and the studies carried out on this aircraft, several aerodynamic optimisations are performed on this geometry with the capability developed in this thesis.

This thesis aims to make advances in the accurate simulation of the ows in and around clustered module plug nozzles. The resulting simulations presented in this thesis are, as far as can be ascertained from available data, the most detailed to date in Europe. A comparison is made with results from other sources for clarication of this point. In the process of producing these solutions, two ow solvers have been developed. NSAXIMB is a general 2D multi-block ow solver,developed by the author, for the axisymmetric, Reynolds averaged Navier-Stokes equations. It was developed to allow simulation of axisymmetic plug nozzle congurations and the investigation of the effects of turbulence modelling on such ows. MERLIN is a general 3D, implicit, multi-block ow solver again for the RANS equations. MERLIN was developed by the Centre for Computational uid Dynamic at Craneld. Signicant input from this work has included a large portion of the structure of the mean ow solver and the extension of the advanced two equation turbulence modelling, incorporated in NSAX- IMB, to three dimensions. Of the turbulence models investigated the zonal models of Menter prove to be most effective in reproducing experimental results. These models out perform a more advanced non-linear eddy viscosity formulation, based on the work of Abid. In an effort to improve solution accuracy, grid adaptation software, based on node redistribution techniques has been developed for use in conjunction with the 3D ow solver. This work is demonstrated in conjunction with a basic test case before application to the clustered module plug nozzle conguration. Results for the complex block topology adopted in the 3D test case are shown to cause the adaptation process to fail. Further, it is shown that such a process may not be generalised for arbitrary topologies.