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An experimental study has been performed at M∞=8.2 and Re∞/cm=93000 to examine: 1. The effect of strakes on the aerodynamic characteristics and performance on slender elliptic cone missile configurations. Some information regarding the shock layer was obtained from schlieren pictures. Detailed flow properties in the shock layer were obtained, for some elliptic cone configurations with and without strakes, using a threedimensional, high resolution, iterative, finite volume parabolized Navier-Stokes solver. Surface flow visualisation, using an oil-dot technique, and pressure measurements were made on one of the models to determine the effect of strakes. Lift, drag and pitching moment characteristics for the elliptic cones with and without strakes were obtained using a three component strain-gauge balance. No gross external flow differences were detected from the schlieren pictures for models tested due to the addition of strakes. Oil-dot visualisation demonstrates that the strakes alter the surface flow characteristics and tended to inhibit the cross-flow. The addition of strakes caused a reduction of pressure on the leeward side and an increase of pressure on the windward side. The strakes produced a significant increase in the lift and drag coefficients, in the incidence range of 0° to 200. The right elliptic cone without strakes with its major axis horizontal exhibits higher lift coefficients than the cone with its major axis vertical. The numerical study predicted the complex flowfield surrounding the right elliptic cone with its major axis horizontal, gave a better understanding of the complicated nature of the flow and good indications of the shock shape and vortex core positions. An estimation model of the aerodynamic forces and moments for the right elliptic cone with and without strakes was developed based on the standard Newtonian theory. The model successfully predicted the experimental trends in the aerodynamic coefficients. 2. The aerodynamic effectiveness of a cylinder flare body at zero incidence under laminar and turbulent boundary layer conditions. Two nose geometries, namely a 10° half-angle sharp cone and a hemisphere, were used. The surface flow over the cylinderflare body was studied using oil-dot and liquid crystal techniques. Some information regarding the shock layer was obtained from schlieren pictures. The effects of entropy layer and boundary layer state on flare effectiveness were deduced from pressure measurements over the cylinder and the flare. The most important difference between the laminar and turbulent boundary layer interaction is that a much smaller angle is necessary to cause laminar separation than that necessary for turbulent separation. The determination of incipient separation is very sensitive to the detection method employed. The existence of a small scale separation bubble can explain the differences in the determination of incipient separation angles if different experimental methods are used. The addition of a hemisphere nose reduces the surface pressure and heat transfer levels on the flare. This is due to loss of reservoir pressure across the bow shock wave. The reduction of flare pressure also reduces the separated flow lengths for the laminar case, whereas for the turbulent case the separated flow lengths are increased. This may be due to the boundary layer along the cylinder body not being fully developed. The effect of Mach shear on the flare pressures distribution has been calculated theoretically. The model predicted the experimental results satisfactorily.

Flapping-wing micro air vehicles, based on insect-like apping, could potentially ll a niche in the current market by o ering the ability to gather information from within buildings. The aerodynamics of insect-like apping are dominated by a large, lift-enhancing leading-edge vortex (LEV). Historically, the cause and structure of this vortex have been the subject of controversy. This thesis is primarily intended to provide insight into the LEV, using computational uid dynamics coupled with validating experiments. The problem is simpli ed by breaking down the complex kinematics involved in insect-like apping and examining only a part of these kinematics; rstly in 2D, before progressing to 3D sweeping wing motions. The thesis includes discussion of published literature in the eld, highlighting gaps and inconsistencies in the current knowledge. Among the contributions of this thesis are: descriptions of the e ects of changing Reynolds number and angle of attack for 2D and 3D ows; clari cation of terminology and phenomenology, particular in the context of 2D ows; and detailed descriptions of the development and structure of the LEV in both 2D and 3D cases, including discussion of Kelvin-Helmholtz instability. The issues of Strouhal number, delayed leading-edge separation, dynamic stall and the Wagner e ect are also considered. Generally, the LEV is shown to be unstable in 2D cases. However, in 3D cases the LEV is seen to be stable, even if Reynolds number is increased. The stability of the LEV is found to be critically dependent on wing aspect ratio.

When analyzing the aerodynamic characteristics of a road vehicle, the flow around the basic body shape is complicated by the presence of the rotating wheels. Even though on most vehicles the wheels are partially shrouded their effect on the flowfield is still considerable. Despite this, very little is understood about the flow around a rotating wheel. This thesis describes the development of a validated steady state Reynolds Averaged Navier-Stokes CFD model to investigate the flow around automobile wheels. As all the previous investigations into the aerodynamic characteristics of wheel flows had been experimental, preliminary computational studies were performed. The basis of these was the 2D circular cylinder. The effects of cylinder rotation and ground proximity were modelled, and strategies for boundary conditions and mesh topology were developed. This work was extended into 3D with the modelling of an isolated wheel, both rotating and stationary. Using existing experimental data for validation, an extensive investigation into the effects of solver numerics, symmetry planes, turbulence models, and the method of turbulent closure was performed. An optimum solver configuration was developed which comprised of the RNG k-E turbulence model with full boundary layer closure. It was accurately predicted that the rotating wheel generates less lift and drag than the equivalent stationary wheel. A number of postulated experimental flow features were captured in the final solutions. Using a parallel experimental study to provide further validation data, the CFD model was extended to incorporate an asymmetric shroud containing a wheelhouse cavity. The influence of the rotation of the wheel, the geometry of the shroud, and the thickness of the stationary groundplane boundary layer were investigated. The rotating wheel now produced more drag than the equivalent stationary wheel. Reductions in wheel drag were found with a reduction in the ride height of the shroud, and with the addition of spoilers to the lower front edge of the shroud. Increasing the stationary groundplane boundary layer thickness also reduced the wheel drag. The effects of these changes on the wheel surface pressure distributions are presented.

Noise is an environmental concern and due to the increasing interest in helicopters as an alternative inter-city transportation, research for more environment friendly helicopters is continuously growing. Building on this demand, this study aims at finding an efficient and accurate noise prediction tool for rotating sources. This study therefore investigates the modelling of noise from rotating sources such as helicopter rotors by addressing noise propagation in both subsonic and transonic/ supersonic regimes. The aim of this research is to explore the field of aeroacoustics prediction for rotor generated noise and to develop a noise prediction tool for sources moving in subsonic or supersonic flow regimes. The aeroacoustics predictions presented have been obtained using a hybrid approach. With such an approach the near field noise generation process is simulated by means of an aerodynamics prediction tool while the noise propagation to the near field is computed by mean of the Ffowcs Williams-Hawkings equation in time domain. For the near-field aerodynamic calculations di erent CFD tools have been exploited. More precisely, three test cases have been analysed. For the first test case of 2D aerofoil-vortex interaction, reproducing the experimental campaign of Lee et al., the near-field is computed via the commercial software Fluent. The unsteady implicit Euler solver with second order discretisation both in space and in time is exploited. This uses the ROE FDS scheme for the fluxes calculation. The same solver is used in the near-field simulations of the third test case, where the analysis of a non-lifting hovering rotor is carried out in delocalised conditions, reproducing the experiments of Purcell on the UH1H model rotor. The second test case analysed is based on the HELISHAPE experimental campaign for the ONERA model rotor in BVI conditions. Two comprehensive codes, from Agusta-Westland and Roma Tre, are used to simulate the complex aeromechanics of the rotor in low speed descent. The noise propagation phase has been performed via the new noise prediction tool developed during this study, named HelicA (for Helic-opter A-coustics). This tool is based on the Emission Surface formulation and exploits a novel root finder and Emission Surface construction algorithms. It can use control surfaces which are in subsonic or transonic/supersonic conditions. Verification and validation processes have been performed on the noise prediction tool before using this code in the aforementioned test cases. These processes are based on the comparison of the tool’s predictions with available analytical and numerical results. The verification and validation cases include sources moving at Mach numbers ranging from MT = 0 to MT > 1. The noise prediction tool is applied to the three aforementioned test cases and the results are in very good agreement with the measurements even for the strong shock delocalisation cases.

Non-linearities within structures often present difficulties when developing algorithms to analyse their dynamic properties. Developing a combined aerodynamic and structural- aeroelastic - code is an example where non-linearities can induce particular characteristics as the presence of aerodynamic non-linearities can compound the complexity of the analysis. Furthermore, when non-linearities occur within actuation devices the impact of coupling control systems with the aeroelastic algorithms - aeroservoelastic - must also be considered. In this work, new methods of analysing aero(servo)elastic systems containing various structural non-linearities are studied. The first technique is used to analyse piecewise linear systems. In this method, aeroelastic equations are recast in a form where the independent variable is the time at which the system reaches a discontinuity, sets of these equations are then combined to form an algebraic set of equations describing a Limit-Cycle Oscillation (LCO). The second technique is applied to more general non-linearities by approximating any discrete non-linearities with trigonometric functions, creating a set of continuous Ordinary Differential Equations (ODEs). For both methods, computational efficiency is achieved by applying numerical continuation to track solution branches. The models analysed in this work are two and three degree-of-freedom aerofoil sections containing non-linearities in their heave, pitch and/or flap freedoms. Four different aerodynamic representations are used, two incompressible codes establish the accuracy of the new methods. The other codes are used to study transonic flows and show good agreement with work based on aeroelastic systems with both linear and non-linear structures. Three different control laws - fixed gain, optimal and adaptive - are also investigated to assess their ability to delay flutter onset and suppress LCOs. Optimal control showed the best overall ability to achieve these aims, although it was found that care must be taken not to destabilise areas below the flutter boundary. Finally, a method of analysing fatigue due to structural non-linearities is investigated. The analysis combines the numerical continuation techniques with the Rainflow method to quantify damage due to simple acceleration-deceleration profiles.

The growing global concern over environmental emissions such as nitrogen oxides and noise sets challenging problems for aero-propulsion engineers. Acoustic waves generated by unsteady combustion not only contribute towards the overall noise transmission, but may also cause thermoacoustic instability in combustors, particularly those designed for low NOx emissions. Combustion noise is generated by unsteady combustion – either by the direct generation of acoustic waves or indirectly by the creation of entropy waves. Entropy waves by themselves are silent, but when accelerated, such as through the combustor exit, they create further acoustic waves known as entropy noise. This thesis aims to study transmitted and reflected combustion noise. Current predictions for noise transmission often assume that the wavelengths of the flow perturbations are large compared to the combustor length, known as the compact assumption. We will develop predictions for finite-length combustors accurate to first-order in frequency. The effect of the interaction between an oscillating shock wave with combustion noise is also studied analytically. The predictions agree with data from numerical simulations. Combustion acoustics reflected at the combustor exit may go on to interfere with the combustion process, setting up a feedback mechanism that may lead to thermoacoustic instability. A modified combustor model is presented to study the effect of dissipation and dispersion of entropy waves on the instability, and it was found that the extent of dissipation or dispersion not only plays a significant role on whether instability occurs, but also determines the dominant frequency of oscillations. Furthermore, analytical and numerical investigations suggest that entropy waves are convected with the flow undissipated, and that modelling improvements may be made to take entropy dispersion into account. The findings in this work provide better tools to understand indirect combustion acoustics and to analyse their importance in both transmitted combustion noise and the thermoacoustic instability experienced by low NOx combustors.

In this thesis a Pascal computer program is presented which calculates a proposed design of parachute from some simple input parameters, of the type specified by a customer to a parachute company. The program reduces by a significant degree time spent by parachute engineers in the preliminary design stages. Parachute design is a process which (in common with much engineering design) can be regarded as consisting of a number of separate calculations. The most suitable method (or methods) for each calculation were selected after a thorough investigation of parachute design techniques. The chosen methods must be sufficiently accurate and readily conform to a computer treatment. The data required by the program have been collected from various sources and are stored in a number of files on a floppy disk. The program is applied to requirements received by a parachute company and results obtained compared with the actual parachutes designed. The program is highly interactive with the user who is able to dispute its selection of values for various parameters. Because the designer can make a rapid and objective choice between a number of methods for various calculations, the existence of this program contributes to his knowledge of the relevance of the parameters involved in, and his understanding of, parachute design. Examples of these techniques are given in the text. Possibilities for expanding and improving the program exist in a number of areas. In some cases the data required for a particular parachute or particular design methods are not available or do not exist. Provision has been made for such data to be included in the program when they are received.

The viscous flow around rectangles defined by afterbody length, B, and cross-stream dimension, A, is investigated through a hybrid discrete vortex method. For uniform flow conditions the effects of varying the side ratio, BIA, the angle of incidence, a, and the Reynolds number, Re, are all considered. Pulsating flow results are reported for rectangular cylinders with B/A values of 0.62, 1.0, 2.0 and 3.0, a B/A=1.0 cylinder inclined at 45° and a circular cylinder. At a fixed Reynolds number, Re=200, the variation of drag coefficient with side ratio shows CD increasing with decreasing B/A. This contrasts with the known result at higher Reynolds number, 104<Re<105, for which a maximum drag occurs close to BIA=0.6. A peak is observed in both the Strouhal number and lift coefficient close to B/A=0.30. This is explained by the afterbody suppression of the shear layer interaction. In the case of the square cylinder, results are presented for the variation of drag coefficient and Strouhal number with Reynolds number, 50<Re5x103. Good agreement with experiment is shown although for Re>500 the calculated Strouhal number is dual valued. The 'lock-in' characteristics under pulsating flow are shown to be highly dependent on body geometry. All the cylinders are shown to exhibit an asymmetric resonant mode within which the shedding frequency is controlled at half the forcing frequency and the mean forces increase. Several different shedding patterns are predicted across this asymmetric synchronisation range. A 'quasi-symmetric' mode is also observed for some cylinders characterised by near wake symmetry and a substantial reduction in mean forces. A pseudo-phase lag is defined which relates a moment of the lift cycle to a moment of the forcing oscillation. This is shown to change across the synchronisation range of each cylinder considered and the change is found to be greater at lower forcing amplitude.

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