Helicopter handling qualities are investigated using inverse simulation as the method of providing state and control information for the appropriate quantitative metrics. The main aim of the work was to develop a more comprehensive and versatile method of quantifying handling qualities levels using the available inverse algorithm "Helin v". Subsequently, the assessment of the helicopter model inherent in Helinv, "Helicopter Generic Simulation", (HGS) for its suitability to handling qualities studies was paramount. Since the Helinv inverse algorithm operates by initially defining a mathematical flight test manoeuvre for the vehicle to "fly", considerable time was given to modelling suitable handling qualities assessment manoeuvres. So-called "attitude quickness" values were then calculated thus providing an initial objective insight into handling qualities level of the vehicle under test. Validation of the tasks formed an integral part of successfully fulfilling the flight test manoeuvre development objective. The influence of the human is captured by the inclusion of a pilot model and the development of a novel method of parameter estimation, supplements the overall objective of modifying Helinv results to achieve potentially more realistic responses and thus correspondingly more realistic handling qualities. A comparative study of two helicopters, one based on the Westland Lynx battlefield/utility type and the other, a hypothetically superior configuration effectively demonstrates the capability of inverse simulation to deliver results adequate for initial handling qualities studies. Several examples are used to illustrate the point. Helinv has been shown to be versatile and efficient and can be used in initial handling qualities studies. The advantages of such a technique are clear when it is seen that actual flight testing, ground based or airborne is extremely costly, as the flight test manoeuvres must be representative of real life, reproducible and of course, as risk free as possible. Many inverse simulation runs and handling qualities calculations have been carried out for different helicopter configurations and manoeuvres thus illustrating the advantages of the technique and fulfilling all the aims mentioned above.

The influence of trailing-edge separation on the dynamic stall characteristics of a typical rotor section is at present unclear. Although previous research has given a fundamental understanding of the unsteady stall process, the variety of aerofoils tested has made it difficult to isolate the effect of trailing-edge separation. Further investigation into this field may be carried out by testing two similar aerofoils which differ only in their trailing-edge separation characteristics. The early part of the work concentrated on the development of a numerical method whereby the theoretical pressure gradient over the trailing-edge upper surface of a given aerofoil may be modified to either enhance or reduce such separation. Since previous work at the University of Glasgow had included a detailed unsteady aerodynamic study of a NACA 23012 aerofoil, this was the appropriate profile for modification. The above technique was applied to this aerofoil with the objective of modifying the geometry in such a manner that would retain the leading-edge pressure distribution whilst forcing an earlier and more gradual trailing-edge separation growth. The subsequently designed aerofoil, designated the NACA 23012(A), was shown to display an enhancement of the trailing-edge separation characteristics via both boundary-layer calculations and oil-flow visualisation tests. On comparison with unsteady data previously collected for the NACA 23012, several systematic methods of estimating the effects of trailing-edge separation on the dynamic stall process are presented. During oscillatory tests the NACA 23012(A) displayed a more stable damping characteristic which was attributed to the enhanced trailing-edge separation producing an earlier pitching-moment break. Based on the analysis of pressure-time histories obtained during ramp tests, it was deduced that a consequence of significant trailing-edge separation was to delay the initiation of the dynamic stall vortex. Detailed analysis of hot-film data led to the conclusion that aerofoils which display a tendency to stall in steady conditions, via separation growth from the trailing-edge, will experience vortex initiation by the breakdown of a thin layer of reversed flow travelling upstream beneath a stable shear layer which remains in close proximity to the aerofoil's surface contour.

Focusing on helicopter aerodynamics, it is known that the aerodynamic performance of the retreating side of a rotor disk is mainly dictated by the stall characteristics of the blade. Stall under dynamic conditions (Dynamic Stall) is the dominant phenomenon encountered on heavily loaded fast-flying rotors, resulting in an extra lift and excessive pitching moments. Dynamic stall (DS) can be idealised as the pitching motion of a finite wing and this is the focus of the present work which includes three main stages. At first, comparisons between available experimental data with CFD simulations were performed for 3D DS cases. This work is the first detailed CFD study of 3D Dynamic Stall and has produced results indicating that DS can be predicted and analysed using CFD. The CFD results were validated against all known experimental investigations. In addition, a comprehensive set of CFD results was generated and used to enhance our understanding of 3D DS. Straight, tapered and swept-tip wings of various aspect ratios were used at a range of Reynolds and Mach numbers and flow conditions. For all cases where experimental data were available effort was put to obtain the original data and process these in exactly the same ways as the CFD results. Special care was put to represent exactly the motion of the lifting surfaces, its geometry and the boundary conditions of the problem. Secondly, the evolution of the Ω-shaped DS vortex observed in experimental works as well as its interaction with the tip vortices were investigated. Both pitching and pitching/rotating blade conditions were considered. Finally, the potential of training a Neural network as a model for DS was assessed in an attempt to reduce the required CPU time for modelling 3D DS. Neural networks have a proven track record in applications involving pattern recognition but so far have seen little application in unsteady aerodynamics. In this work, two different NN models were developed and assessed in a variety of conditions involving DS. Both experimental and CFD data were used during these investigations. The dependence of the quality of the predictions of the NN on the choice of the training data was then assessed and thoughts towards the correct strategy behind this choice were laid out.

The numerical solution of the two-dimensional inviscid Euler flow equations is given. The unstructured mesh is generated by the advancing front technique. A cell-centred upwind finite volume method has been adopted to discretize the Euler equations. Both explicit and point implicit time stepping algorithms are derived. The flux calculation using Roe's and Osher's approximate Riemann solvers are studied. It is shown that both the Roe and Osher's schemes produce an accurate representation of discontinuities (e.g. shock wave). It is also shown that better convergence performance has been achieved by the point implicit scheme than that by the explicit scheme. Validations have been done for subsonic and transonic flow over airfoils, supersonic flow past a compression corner and hypersonic flow past cylinder and blunt body geometries. An adaptive remeshing procedure is also applied to the numerical solution with the objective of getting improved results. The issue of high order reconstruction on unstructured grids has been discussed. The methodology of the Taylor series expansion is adopted. The calculation of the gradient at a reference point is carried out by the use of either the Green-Gauss integral formula or the least-square methods. Some recently developed limiter construction methods have been used and their performance has been demonstrated using the test example of the transonic flow over a RAE 2822 airfoil. It has been shown that similar pressure distributions are obtained by all limiters except for shock wave regions where the limiter is active. The convergence problem is illustrated by the mid-mod type limiter. It seems only the Venkatakrishnan limiter provides improved convergence. Other limiters do not appear to work as well as that shown in their original publications. Also the convergence history given by the least-square method appears better than that by the Green-Gauss method in the test.

Helicopter agility in nap-of-the-earth flight is widely recognised to be of great importance. Despite this. a general method of quantifying agility does not exist. All previous attempts to quantify agility have been restricted either to flight tests or to simple kinematic modelling - both with obvious disadvantages. A method of quantifying helicopter inherent agility. the agility of the configuration independent of the pilot. utilising inverse solutions of the equations of motion has been developed. A value for the inherent agility of a helicopter is given by studying its performance over a series of standard manoeuvres. The manoeuvres used represent typical tasks undertaken by the configuration under study. The combination of these tasks represent the helicopter's operational role. The helicopter's performance over these standard manoeuvres is found by using an inverse solution of the equations of motion - calculation of the control. and resulting state. time histories needed to fly a given flight path. A six degrees of freedom non-linear mathematical model is used to simulate single main and tail rotor helicopter flight dynamics. The helicopter's performance over each manoeuvre is rated by a quadratic performance function of the state and control variables. The performance function is weighted in such a manner as to penalise undesirably large displacements in the state and control variables of particular importance to that manoeuvre (e.g. large nose down attitude changes in accelerated flight are heavily penalised). An Agility Rating is awarded to a helicopter on the basis of its performance over a wide range of similar manoeuvres. a measure of total inherent agility being a function of the agility ratings for all the manoeuvres relevant to the helicopter's role. The method is illustrated by applying it to two agility studies. Firstly. it is used to show how an optimum tailplane area can be calculated for manoeuvres in the longitudinal plane. Then an "Advanced Rotor Helicopter" is compared with a contemporary battlefield helicopter.

To explore the influence of wind tunnel test facilities on delta wing aerodynamics, the interference has been separated into two distinct types, wall interference and support structure interference. The wall interference effects have been split into three further components, tunnel blockage, side wall interference, and roof and floor interference. Splitting the tunnel influence in this way allows us to determine the most detrimental interference effects, thus allowing the wind tunnel engineer to design experiments accordingly. Euler and more realistic RANS simulations of tunnel interference have been conducted. To reduce the question of grid dependence when comparing solutions, a common "farfield grid" was created and tunnel grids were extracted. Before doing RANS simulations an analysis of various turbulence models was conducted. It was found that turbulence models have difficulty in predicting turbulence levels in leading edge vortices. As such modifications have been applied to the models which improve predictions. Despite vortex breakdown being widely regarded as an inviscid phenomenon, dependence on turbulence modelling has been exhibited. This is due to the vortex properties being altered with turbulent diffusion of vorticity when turbulence levels are too high. Both 1- and 2-equation models were assessed and it was concluded that a modified 2-equation k-w model was the most suitable of the models available (when compared against experimental results), and was therefore used in all subsequent simulations. From both Euler and RANS simulations it has been concluded that the effect of sidewall proximity significantly promotes vortex breakdown. Side wall induced velocity components increase the mean effective incidence of the wing, the helix angle and the strength of the vortices. The combination of these effects promotes vortex breakdown. Roof and floor proximity has little effect on vortex breakdown as does the frontal area blockage. Pitching simulations have shown that the promotion of vortex breakdown is not consistent on both the upstroke and downstroke. Break-down was observed to be prompted furthest at the higher incidence of the upstroke and on the downstroke. This highlights the dependency of tunnel interference on vortex strength.

For the simulation of control surface buzz accurate prediction of the shock location and the chock strength is essential and this is currently achieved using Euler and RANS based CFD analysis. To calculate the motion of the control surface only the flap rotation mode needs to be modelled. In the current work the CFD solver is coupled with a modal based FEM solver. The multi-level hierarchical blending transformation methodology is applied for the aeroelastic analyses of complex geometries. The methodology is used for the treatment of blended control surfaces and the effect of the blending on the aero-structural response is measured. Forced clap oscillations of a Supersonic Transport (SST) configuration are simulated and the dynamic deformation of the wing and the unsteady pressure due to the forced oscillations are validated against experiments. Transonic buzz on a trailing edge flap is investigated on the Supersonic Transport configuration using the RANS and the Euler equations. Characteristics associated with buzz instability are reproduced computationally and the effect of the flap on the wing flutter is measured. Finally, aeroelastic simulations are performed on the Hawk aircraft. The combat flap configuration of the Hawk aircraft is investigated using CFD and the effect of the flap on wing flutter is assessed. The aeroelastic response of the rudder at supersonic freestream Mach numbers is studied. The importance of aerodynamic interference on the aeroelastic behaviour is assessed.

After introducing the main features of the aerodynamics of wind turbines, a review of key theoretical studies and aerodynamic modelling methods provides the opportunity to focus on predictive methods and the main technical challenges associated with the aerodynamics of HAWTs. The basic aerodynamic method adopted in this study, a classic Blade Element Momentum theory model, BEM, is described next and its extension to yawed flow is detailed for completeness. Analysis then focuses on how stall delay due to three-dimensional effects can be predicted on a HAWT. Implementation of a semi-empirical stall delay model shows sensitivity to blade geometry but no dependency on wind velocity or rotational speed. This seems to be physically incorrect and suggests that a deeper understanding of 3-D effects is still needed if better algorithms are to be developed. The work then examines the onset of dynamic stall. A 2-D semi-empirical correlation of vortex stall onset, developed previously at Glasgow University, is implemented and validated through available field data from the NREL turbine Phases II and IV. The comparison of measured and predicted locations of dynamic stall onset highlights some interesting features of the three-dimensionality of the process; after the local inception, earlier dynamic stall appears to be triggered in adjacent stations. An attempt to study how 3-D stall delay interacts with the onset of dynamic stall, shows that stall delay appears not to influence the inception of dynamic stall in the way it does static stall. Moreover, the firsts signs of dynamic stall onset are generally best characterised by the correlation when it assumes locally 2-D flow. This is a significant result, as it demonstrates that the earliest signs of dynamic stall onset on wind turbines can be correctly predicted using 2-D tools. A closer examination of the discrepancies between the predictions and measurements has highlighted the particular aerodynamic characteristics of the S809 aerofoil, utilised as the blade section of the NREL turbines. The unusual stalling characteristics of this aerofoil bring into question the significance of the static stall angle in relation to dynamic stall. It is show that other features of the static behaviour may provide a more appropriate link to dynamic stall for some aerofoils. Finally, the phenomenon of tower shadow on a downwind turbine is studied. Unaveraged pressure measurements and integrated normal force coefficients from tests conducted at Glasgow University are analysed. The analysis highlights many interesting features of the tower shadow response. In particular, as the blade enters the tower shadow region, there is a rapid reduction in normal force due to the tower wake velocity deficit. As the blade leaves the tower shadow, the recovery is consistently slower and more progressive and apparently extends further than the edge of the velocity deficit region. These observations are then used in a examination of tower shadow modelling. A steady model, based on a cosine shaped velocity deficit is evaluated by comparison with the wind tunnel measurements. Unfortunately neither the phase nor the intensity of the response is adequately captured. This leads to the implementation of a new model, based on classic unsteady thin aerofoil theory that accounts for the aerofoil wake induced velocities. The unsteady model captures, in a satisfactory manner, the global response of the blade through the tower shadow region, with a negligible computational cost.

Possible helicopter flight mechanics benefits associated with the use of an actively controlled horizontal tailplane are identified, influencing the areas of agility and manoeuvrability. In both cases, control strategies are postulated and implemented by means of control laws. They are then used with mathematical descriptions of the helicopter in digital computer simulations of manoeuvres to quantitfy the benefits. In the field of helicopter agility, use of a relatively small horizontal tailplane is shown to enhance agility, relative to the helicopter with a fixed tailplane. Popup maneouvres to SOm can be flown up to 7% faster with the active tailplane; alternatively, geometrically tighter manoeuvres can be flown to the extent of reducing manoeuvre distance by up to 10%. The control law moves the tailplane proportionally with the contributions of the three rotor controls and helicopter pitch rate to the longtitudinal component of hub moment. It is however suggested that a tailplane control law based on functions of pitch attitude would be applicable to a wider range of manoeuvres than the popups simulated. Helicopter manoeuvrability is enhanced by using the tailplane to decouple the pitch attitude from the flight path. The benefits are demonstrated by simulation of the acquisition and tracking of an airborne target. For a helicopter with the conventional pattern of control, significant changes in flight path result when the target is tracked with fuselage pointing; by comparison, the helicopter with a decoupled flight path and attitude controller changes flight path and speed by a negligible amount. It is suggested that this mode of control may be more generally applicable to control of the helicopter in that it mitigates the speed/flight path/attitude compromise the pilot faces in flying his aircraft, or the possibly large hub moments when accelerating or decellerating. The philosophy behind the use of the active tailplane differs from that of contemporary applications of moveable tailplanes in that it is an integrated element of the flight control system endowing (in its own right) control capabilities on the helicopter that are otherwise precluded by configuration. The addition of this extra control demands active control technology for several reasons: the applications require full control authority; the control laws are multivariable and change with speed; and the cockpit control setup would have to be simplified to the extent of the radical changes facilitated by active control technology.

In this thesis, a selection of numerical optimisation methods were developed for application to realistic solar sail heliocentric trajectory optimisation problems. A Non-Linear Programming method based on Sequential Quadratic Programming was developed, with the sail controls parameterised in time. This method was hybridised with Genetic Algorithms or locally-optimal analytical control laws to generate an initial guess, where required. The goal of this thesis is to create a detailed catalogue of trajectories to a broad range of heliocentric targets, subject to realistic constraints on trip-time, sail performance, and thermal-limited solar approach. This thesis illustrates the wide range of targets in the solar system that can be reached with solar sailing. In addition, the trajectory problems for which solar sailing is not attractive are also identified. Trajectory analysis of sample return missions to the terrestrial planets, Mars, Venus and Mercury, has been conducted. Extensive depmiure date scans were perfOlmed, where it was found that there are minima and maxima in trip-time, separated by discontinuities, providing effective launch windows. Roundtrip optimal launch dates were identified, after combining outbound and return departure date scans. For Mercury rendezvous, the application of positive launch excess velocity and a Venus gravity assist was investigated, where a small trip-time saving can be made. Trajectories to rendezvous with the Short Period Comet Wirtanen have been optimised, where it was found that a significant reduction in trip time and launch mass could have been realised, relative to a conventional mission. An investigation of using higher performance sails to flyby Long Period Comets has also been conducted, to demonstrate that solar sailing could be used to reach newly discovered comets soon after first discovery, such as the previous Hale-Bopp apparition. It is also shown that solar sailing could be used, instead of solar electric propulsion, to rendezvous with two Main-Belt asteroids, with a reduction in launch mass. The open ended nature of solar sailing was used to show that rendezvous with two further asteroids is also possible. It is also shown that a three-phase trajectory concept, utilising an inclination cranking manoeuvre, could be used to return a sample from a high inclination Near-Earth Asteroid, that would be essentially impossible to reach using conventional propulsion. It is demonstrated that flyby missions to the outer planets, such as Pluto are feasible in reasonable timescales using a solar photonic assist concept. However, due to the faint solar radiation pressure at Jupiter, only flyby missions are practical to the Jovian system with solar sails. An extensive trade-off between launch hyperbolic excess energy, Jupiter arrival velocity, hip-time, and the number of photonic assist loops has been conducted. By contrast, solar sailing appears to be the only feasible option for missions to the Heliopause at 200 AD. Heliopause trajectory analysis included investigation of the number of loops, and the effect of thermallyconstrained closest solar approach on escape velocity and trip-time. It was found that, in order to reach the Heliopause in 25 years, a solar sail of characteristic acceleration of order 1.5 mm S-2 would be required, executing a thermally constrained solar photonic assist at 0.25 AD. Investigation of the effect of positive launch energy is also conducted for Heliopause trajectOlies. A key near-term mission application for solar sails is a Solar Polar Orbiter. Trajectory analysis has revealed that a solar sail transfer to a true solar polar orbit, Earth resonant at 0.48 AU, in 5 years would require a characteristic acceleration of 0.42 nun S-2. In the course of the parametric analysis, two-phase and three-phase scenarios were investigated, with an assessment of the effect of spiralling down to a close cranking orbit radius from positive launch excess energy. Finally, new transfers to exotic, displaced Non-Keplerian Orbits have been optimised for a range of final orbit dimensions among one family of these unique orbits. For lower performance sails, transfers to artificial Lagrange points have been optimised, in the context of the Geostorm and Polar Observer missions.