This research will investigate various swept wing models, designing the mechanism for their pitching motion and control, designing wind tunnel implementation, and performing data measurements and analysis using particle image velocimetry. A NACA0012 section with an aspect ratio of AR = 4, free stream velocity of U∞=34 m/s, and Reynolds Number is Rec=2x105. Swept airfoils of Λ=0°, 15°, and 30° will be pitched sinusoidally between an AoA of 4°and 22°, at a reduced frequency of k=πfc/U∞=0.2. Higher sweep angles developing arch-type vortices interact with wing tip flow and abrupt tip stall is observed. Lower sweep angles possessed defined leading edge vortices persist near the tip after lift has collapsed at mid span. Stall angle was delayed during dynamic motion of the wing as well as the presence of arch and ring type vortices increased with sweep angle and contributed to flow reattachment along the top surface of the wing.

This thesis is focused on the aerodynamic analysis of the competition paraglider wing. Drags of the particular wing parts are divided into chapters. The aim was to get a grasp of sizes of the individual components drags in relation to the entire assembly. In the first instance, a 2D profile and then the entire configuration of the 3D wing was analyzed. After the evaluation, some power reserves were detected in an airfoil and so the airfoil shape was optimized. After the optimization of the individual components, the CFD calculation was used again. At the end, geometry changes were evaluated.

The purpose of this research is to evaluate certain performance characteristics of a morphing<br />wing system that uses Macro Fiber Composites (MFC) to create camber change. This<br />thesis can be broken into two major sections. The first half compares a few current MFC<br />airfoil designs to each other and to a conventional servomechanism (servo) airfoil. Their<br />performance was measured in terms of lift and drag in a 2-D wind tunnel. The results<br />showed MFC airfoils were effective but limited by aeroelasticity compared to the servo. In<br />addition, a morphed airfoil and a flapped airfoil were rapid prototyped and tested to isolate<br />the effects of discontinuity. The continuous morphed airfoil produced more lift with less<br />drag.<br />The second half of this thesis work focused on determining the ideal MFC configurations for<br />a thin wing application. Simulations were run on a thin wing with embedded MFCs such<br />that the whole wing morphed. Finite element and vortex lattice models were used to predict<br />deflections and rolling moment coefficients. Different configuration parameters were then<br />varied to quantify their effect. The comparisons included MFC location, number of MFCs,<br />material substrate, and wing thickness. A prototype wing was then built and flight tested.<br />While the simulations overestimated the wing deflection, the flight results illustrated the<br />complexity and variability associated with the MFC morphing system. The rolling moment<br />coefficients from flight were consistent with the simulation given the differences in deflection. / Master of Science

The study of designing stable aircraft has been widespread and ongoing since the early days of Orville and Wilbur Wright and their famous Wright Flyer airplane. All aircraft as they fly through the air are subject to minor changes in the forces acting on them. The field of aircraft stability seeks to understand and predict how aircraft will respond to these changes in forces and to design aircraft such that when these forces change the aircraft remains stable. The mathematical equations used to predict aircraft stability rely on knowledge of the location of the aerodynamic center, the point through which aerodynamic forces act on an aircraft. The aerodynamic center of an aircraft is a function of the aerodynamic centers of each individual wing, and the aerodynamic center of each wing is a function of the aerodynamic centers of the individual airfoils from which the wing is made. The ability to more accurately predict the location of the airfoil aerodynamic center corresponds directly to an increase in the accuracy of aircraft stability calculations. The Aerolab at Utah State University has develop new analytic mathematical expressions to describe the location of the airfoil aerodynamic center. These new expressions do not suffer from any of the restrictions, or approximations found in traditional methods, and therefore result in more accurate predictions of airfoil aerodynamic centers and by extension, more accurate aircraft stability predictions.

The research presented in this thesis examines the use of an estimated "actuator work" value as a performance parameter for the comparison of various aerodynamic control device configurations. This estimated "actuator work," or practical work as it will be referred to as in this thesis, is based on the aerodynamic and structural resistance to a control surface deflection. It is meant to represent the actuator energy cost required to deflect a general configuration of conventional or unconventional control surface. Thin airfoil theory is used to predict the aerodynamic load distribution required for this work calculation. The details of applying thin airfoil theory to many different types of control surface arrangements are presented. Convenient equations for the aerodynamic load distributions and aerodynamic coefficients are obtained. Using the developed practical work equations, and considering only the aerodynamic load component, the practical work required for a given change in lift is compared between different control surface arrangements. For single control surface cases, it is found that a quadratic (morphing) trailing edge flap requires less practical work than a linear flap of the same size. As the angle of attack at which the change in lift occurs increases, the benefit of the quadratic flap becomes greater. For multiple control surface cases, it is necessary to determine the set of control deflections that require the minimum practical work for a given change in lift. For small values of the initial angle of attack, it is found that a two-segment quadratic trailing edge flap (MTE) requires more work than a two-segment linear flap (TETAB). But, above a small value of angle of attack, the MTE case becomes superior to the TETAB case. Similar results are found when a 1-DOF static aeroelastic model is included in the calculation. The minimum work control deflections for the aeroelastic cases are shown to be strongly dependent on the dynamic pressure. / Master of Science

No description available.

A neural network prediction method has been developed to compute self-noise of airfoils typically used in wind turbines. The neural networks were trained using experimental data corresponding to tests of several different airfoils over a range of flow conditions. The experimental data corresponds to the NACA 0012, Delft DU96, Sandia S831, S822 and S834, Fx63-137, SG6043 and SD-2030 airfoils. The chord of these airfoils range from 0.025 to 0.91 m and they were tested at Reynolds numbers of up to 3.8 million and angle of attack up to 15° depending on the airfoil. Using experimental data corresponding to different airfoils provides to the neural network the capacity to take into account the geometry of the airfoils in the predictions.geometry of the airfoils in the predictions. The input parameters to the network are the flow speed, chord length, effective angle of attack and parameters describing the geometrical shape of the airfoil. In addition, boundary layer displacement thickness was used for some models. The parameters used for taking into account the airfoil's geometry are based on a conformal mapping method or a polynomial approximation. The output of the neural network is given by sound pressure level in 1/3rd octave bands for nine frequencies ranging from 630 to 4000 Hz. The present work constitutes an application of neural networks to aeroacoustics. The main objective was to assess the potential of using neural networks to model airfoil noise. Therefore, this work is focused in the modeling of the problem, and no mathematical analyses about neural networks are intended. To this end, several models were investigated both in terms of the configuration and training approach. The performance of the networks was evaluated for a range of flow conditions. The neural network technique was first investigated for the NACA 0012 airfoil only. For this case, the geometry of the airfoil was not incorporated as input into the model. The neural network approach was then extended to account for airfoils of any geometry by including data from all airfoils in the training. The results show that the neural networks are capable of predicting the airfoils self-noise reasonably well for most of the flow conditions. The broadband noise due to the turbulent boundary layer interacting with the trailing edge is estimated very well. The tonal vortex shedding noise due to laminar boundary layer-trailing edge interaction is not predicted as well, most likely due to the limited data available for this noise source. In summary, the research here demonstrated the potential of the neural network as a tool to predict noise from typical wind turbine airfoils. / Master of Science

The automation of airfoil design process is an ongoing effort within the field of turbo-machinery design, with significant focus on developing new reliable and consistent methods that can meet the needs of the engineers. A wide variety of approaches has been in use for inverse airfoil design process which benefit from theoretical inverse design, statistical methods, empirical discoveries and many other ways to solve the design problem. This thesis work also develops a tool in Python to be used in airfoil aerodynamic design process that is simple, fast and accurate enough for initial design of turbo-machinery blades with focus on turbine airfoils used for operation in aircraft engines. To convey the decision-making process during development a simplified case is presented. The underlying considerations are discussed. Other available methods in the literature used for similar problems, are also evaluated and compared to demonstrate the advantages and limitations of the methods used within the tool. The inverse design problem is formulated as a multi-objective optimization problem to handle various different objectives that are relevant for aerodynamic design of turbo-machinery airfoils. Test runs are made and the results are discussed to assess how robust the tool is and how the current capabilities can be modified or extended. After the development process, the tool is verified to be a suitable option for real-life design optimization tasks and can be used as a building block for a much more comprehensive tool that may be developed in the future. / Automatisering av processen för design av vingprofiler kräver fortlöpande insatser inom området turbomaskindesign, med stort fokus på att utveckla nya tillförlitliga och konsekventa metoder som kan tillgodose ingenjörernas behov. Ett stort antal olika tillvägagångssätt har provats för omvänd design av vingprofiler såsom teoretisk invers design, statistiska metoder, empiriska upptäckter och många andra sätt att lösa designproblemet. Detta avhandlingsarbete är också ett lyckat försök att utveckla ett verktyg i Python som ska användas i den aerodynamiska designprocessen; det är enkelt, snabbt och noggrant för den initiala designen av turbomaskinblad med fokus på turbinblad som för användning i flygmotorer. För att förmedla beslutsprocessen under utvecklingen presenteras ett förenklat fall. De underliggande övervägandena diskuteras. Andra tillgängliga metoder i litteraturen som används för liknande problem utvärderas och jämförs för att visa fördelarna och begränsningarna med de metoder som används i verktyget. Det omvända designproblemet formuleras som ett multi-objektivt optimeringsproblem för att hantera olika mål som är relevanta för aerodynamisk design av turbomaskiner. Testkörningar görs och resultaten diskuteras för att bedöma hur robust verktyget är och hur de nuvarande funktionerna kan modifieras eller utökas. Efter utvecklingsprocessen verifieras verktyget som ett lämpligt alternativ för verkliga designoptimeringsuppgifter och kan användas som en byggsten för ett mycket mer omfattande verktyg som kan utvecklas i framtiden.

This thesis aims to investigate and reduce the aerodynamic noise source known as trailing edge noise, or airfoil self-noise, by using passive flow control techniques. Airfoil self-noise is produced when a turbulent boundary layer generated on an airfoil surface is scattered by the airfoil’s trailing edge. The investigation is of experimental nature, conducted in the aeroacoustic as well as aerodynamic wind tunnel facilities at Brunel University London and the Institute of Sound and Vibration (ISVR) at the University of Southampton. The research is relevant for any application in which airfoil blades encounter a smooth non-turbulent inflow and hence where trailing edge noise is a dominant noise source. Potential applications can therefore be fan or rotor blades in aero-engines, wind turbine blades or industrial cooling fans. The approach taken for the reduction of trailing edge noise utilises passive flow control techniques through the use of trailing edge serrations and the additional support of porous materials. Both of the aforementioned are inspired by the owl’s silent flight due to its unique wing structure. The research presented here can be divided in three parts: The first part comprises an extensive assessment of the performance of non-flat plate trailing edge serrations for airfoil broadband noise and their aerodynamic performance in terms of lift and drag. It is found that serrations can realistically achieve noteworthy broadband airfoil self-noise reductions, however due to the fact that non-flat plate serrations are directly cut into the airfoil body, the blunt sections in the serration root produce an additional noise source of vortex shedding tonal noise. The second part investigates the two flow mechanisms involved. Regarding the mechanism responsible for broadband noise and the subsequent reductions by the serration geometry, the turbulent boundary layer structures are studied in depth on a serrated trailing edge of a flat plate. Experimental techniques such as hot wire anemometry, liquid crystal flow visualisation, unsteady surface pressure measurements and noise measurements are used. A redistribution of the momentum and turbulent energy near the sawtooth tip and side edges appears to reduce the trailing edge noise scattering-efficiency of the hydrodynamic pressure waves. For the study of the flow mechanism responsible for the vortex shedding tonal noise increase, noise and velocity measurements along with flow visualisation techniques are used for the identification and further understanding of this noise source. A highly three-dimensional wake-flow could be identified in the wake past the serration gap, which differs from the longitudinal vortices shed from a straight blunt serration root. The third part presents the concept of poro-serrated trailing edges as a novel method to substantially improve the overall noise performance of the non-flat plate trailing edge serration type. The use of porous metal foams or thin brush bundles which fill the interstices between adjacent members of the sawtooth can completely suppress the bluntness-induced vortex shedding noise. Most importantly a turbulent broadband noise reduction of up to 7 dB can be achieved without compromising the aerodynamic performances in lift and drag. The new serrated trailing edges do not cause any noise increase throughout the frequency range investigated here. Through noise and velocity measurements near the trailing edge of an airfoil, the reduction of the broadband noise is found to be primarily caused by the sawtooth geometry. The new serrated trailing edges have the potential to improve the industrial worthiness of the serration technology in achieving low noise radiation.

A stand-alone genetic algorithm (GA) and an surrogate-based optimization (SBO) combined with a GA were compared for accuracy and performance. Comparisons took place using the Ackley Function and Rastrigin's Function, two functions with multiple local maxima and minima that could cause problems for more traditional optimization methods, such as a gradient-based method. The GA and SBO with GA were applied to the functions through a fortran interface and it was found that the SBO could use the same number of function evaluations as the GA and achieve at least 5 orders of magnitude greater accuracy through the use of surrogate evaluations. The two optimization methods were used in conjunction with computational fluid dy- namics (CFD) analysis to optimize the shape of a bumpy airfoil section. Results of opti- mization showed that the use of an SBO can save up to 553 hours of CPU time on 196 cores when compared to the GA through the use of surrogate evaluations. Results also show the SBO can achieve greater accuracy than the GA in a shorter amount of time, and the SBO can reduce the negative effects of noise in the simulation data while the GA cannot.