Dissertations / Theses on the topic 'Aerodynamic control surfaces'

Thesis (MScEng (Electrical and Electronic Engineering))--University of Stellenbosch, 2010.<br>ENGLISH ABSTRACT: This thesis describes the first project relating to the Control Surfaces in Confined Spaces (CoSICS) project at Stellenbosch University. The aim of CoSICS project is to reduce the size of control surface actuators, and this thesis considers the aileron system of commercial aircraft such as the Airbus A320 and A330. Specifically the project aims to reduce the aileron hinge moment, as this will result in smaller actuators. Possible methods are discussed where aerodynamic forces are used to reduce the aileron hinge moment through the use of a wing-aileron-tab configuration. In order to examine the use of the configuration, first order aerodynamic modelling is performed using two-dimensional thin-aerofoil theory, which is also extended to a basic three-dimensional approximation. To determine the maximum reduction in hinge moment several optimisations are performed where only the tab chord length is varied, both tab and aileron chord lengths are varied, and finally the tab chord length and aileron span are varied. The optimisation methods used, namely the gradient-based sequential quadratic programming (SQP) and a real-encoded genetic algorithm (REGA) are discussed in detail and include general implementations which are then applied to the problem. The optimisations performed are dual-layered where optimal deflection angles are determined as well as the optimal geometry. The results of the optimisation are tested using a roll manoeuvre in a specially developed Simulink simulation environment for this purpose. The study produces results where new hinge moment values are an order of magnitude smaller than those of the old configuration, while maintaining suitable lift and rolling moment coefficients. The optimisation and simulation infrastructure developed in this thesis provides a platform for higher-fidelity models and components being developed in future work to provide higher fidelity results.<br>AFRIKAANSE OPSOMMING: Hierdie tesis beskryf die eerste projek in die Control Surfaces in Confined Spaces-projek1 (CoSICS-projek) uitgevoer by die Universiteit Stellenbosch. Die doel van die COSICs-projek is om die grootte van beheervlak aktueerders te minimeer en hierdie tesis handel oor die aileron stelsel van kommersiële vliegtuie soos die Airbus A320 en A330. Die doel van hierdie tesisis om die skarnier draaimoment van die aileron te minimeer deur aërodinamiese kragte in te span in ’n vlerk-aileron-hulpvlak konfigurasie. Eerste-orde aërodinamiese modelle is afgelei met behulp van twee-dimensionele dunvlerkteorie en is gebruik om die konfigurasie te analiseer. ’n Eerste orde drie-dimensionele benadering is ook ontwikkel. Om die maksimum vermindering in die skarnier draaimoment te bepaal, is verskeie optimerings uitgevoer waar eers die hulpvlak se koordlengte gevarieer word, daarna beide die aileron en hulp-vlak se koordlengtes en laastens die hulp-vlak se koordlengte en wydte. Die twee optimerings metodes wat gebruik is, nl. ’n sekwensiële kwadratiese programmerings (SKP) tegniek, en ’n reële getal-geënkodeerde genetiese algoritme (RGGA), word bespreek en ontwikkel voor hulle toegepas word op die probleem. Twee-vlak optimerings word uitgevoer waar beide die optimale defleksiehoeke en die optimale geometrie bepaal word. Die resultate van die optimering word daarna getoets deur middel van ’n rol maneuver wat uitgevoer word in ’n Simulink simulasie omgewing wat daarvoor geskep is. Hierdie studie lei tot goeie resultate met skarnier draaimoment waardes ’n ordegrootte kleiner as dié van die vorige stelsel, terwyl goeie waardes van rol-moment en verheffingskrag koëffisiënte behou word. Die optimering en simulasie infrastruktuur wat hier ontwikkel word verskaf ’n platform vir meer akkurate modelle en komponente wat ontwikkel word in toekomstige projekte om meer akkurate resultate te lewer.

With the advent of steerable, ram air parafoil canopies, aerial payload delivery has become a viable alternative for situations involving remote or undeveloped areas, hostile environments, or otherwise inaccessible locations. Autonomously guided systems utilizing such steerable, ram air canopies are typically controlled by symmetric and asymmetric deflection of the canopy trailing edge. Although these systems have demonstrated substantial improvement in landing accuracy over similarly sized unguided systems, their low number of available control channels and limited ability to alter vehicle glide slope during flight makes them highly susceptible to atmospheric gusts and other unknown conditions near the target area. This research aims to improve landing accuracy in such adverse conditions by replacing the standard trailing edge deflection control mechanism in favor of upper surface canopy spoilers. These spoilers operate by opening several spanwise slits in the upper surface of the parafoil canopy thus forming a virtual spoiler from the stream of expelled pressurized air. In particular, estimation of steady-state vehicle flight characteristics in response to different symmetric and asymmetric spoiler openings was determined for two different small-scale test vehicles. Additionally, improvements in autonomous landing accuracy using upper surface spoilers in a combined lateral and longitudinal control scheme was investigated computationally using a high fidelity, 6-DOF dynamic model of the test vehicle and further validated in actual flight experiments with good results. Lastly, a novel in-canopy bleed air actuation system suitable for large-scale parafoil aircraft was designed, fabricated, and flight-tested. The in-canopy system consists of several small, specifically designed wireless winch actuators mounted entirely inside the parafoil canopy. Each in-canopy actuator is capable of opening one or more upper surface canopy spoilers via a unique internal rigging structure. This system demonstrates not only the applicability of bleed air spoiler control for large-scale autonomous parafoil and payload aircraft, but also provides the potential for significant savings in size, weight, and cost of the required actuation hardware for currently fielded systems.

This thesis aims at presenting the Discrete Vortex Method (DVM) as a tool to determine the flow field and associated wind loads over structures. Two structures are considered: the first is a surface-mounted prism and is used to simulate wind loads over low-rise structures. The second is a bridge section with attached flaps that can be oriented to vary the moment coefficient. Advantages and disadvantages of using DVM for these applications are discussed. For the surface-mounted prism, the results show that the developed code correctly predicts the flow separation around the corners. As for the surface pressures, it is concluded that parallel processing, which could be easily implemented for DVM, should be used to correctly predict surface pressures and their variations. This is due to the required slow time advancement of the computations. The results on attaching flaps to bridge sections yield required orientations to minimize moments under different angles of attack.<br>Master of Science

In the past, 2D CFD analysis of Circulation Control technology have shown poor comparison with experimental results. In Circulation Control experiments, typical results show a relationship between lift coefficient, CL, vs blowing momentum coefficient, Cμ. CFD analysis tend to over-predict values of CL due to gridding issues and/or turbulence model selection. This thesis attempted to address both issues by performing Richardson’s Extrpolation method to determine an acceptable mesh size and by using FLUENT’s 2-equation turbulence models. The experimental results and CAD geometry were obtained from Georgia Tech Research Institute for comparison with the CFD analysis. The study showed that 3D CFD analysis of circulation control showed similar results of over-predicting CL, which can also be attributed to gridding issues and turbulence model selection. When compared to the experimental results, the k − ω turbulence model produced the lowest errors in CL of approximately 15-17%. The other turbulence models produced errors within 5% of k − ω. A fully unstructured volume mesh with prismatic cells on the surfaces was used as the grid. The CCW con- figuration was analyzed with and without wind tunnel walls present, which produced errors of 20% and 15% in CL, respectively, when compared to experimental results. Despite the large errors in CL, CFD was able to capture the trend of increasing CL as Cμ was increased. Results reported in this thesis can be further calibrated to allow the CFD model to be used as a predictive tool for other CCW applications.

Thesis (MScEng)-- Stellenbosch University, 2013.<br>ENGLISH ABSTRACT: In this thesis the design, analysis, implementation, and verification of a fault-tolerant unmanned aerial vehicle (UAV) flight control system which is robust to structural damage causing the natural flight dynamics of the vehicle to become asymmetric, is presented. The main purpose of the robust control architecture is to maintain flight stability after damage has occurred. The control system must be able to handle an abrupt change from an undamaged to a damaged state, and must also not depend on explicit knowledge of the damage. A robust control approach is therefore preferred above an adaptive control approach. As a secondary objective, the system must provide robust flight performance to ensure adequate response times and acceptable transients’ behaviour, both in normal flight, and after damage has occurred. An asymmetric six degrees of freedom equations of motion model is derived. The model accounts for the changes in the aerodynamic model of the aircraft as well as changes in the centre of gravity location. Vortex lattice techniques are used to determine the aerodynamic coefficients of the aircraft for damage to the main wing resulting in 0% to 40% spanwise lifting surface loss. A sequential quadratic programming optimisation algorithm is applied to the force and moment equations to find the trim flight state and actuator deflections of the asymmetric aircraft for constant airspeed and altitude. The trim flight state can be further constrained to force zero bank angle, zero sideslip angle or a desired relative weighting of nonzero bank angle and nonzero sideslip angle. The calculated trim actuator deflections are compared to the physical deflection limits to determine the feasibility of maintaining trim flight for different percentages of wing loss. Assuming that a valid trim condition exists, the relative stability of the aircraft’s natural modes is analysed as a function of percentage wing loss by tracing the locus of the open-loop poles. An acceleration-based flight control architecture is designed and implemented, and the robustness of the flight control stability and performance is analysed as a function of percentage wing loss. The robustness and performance of the flight control system is verified with a nonlinear simulation for spanwise wing loss from 0 to 40%. Practical flight tests are performed to verify the robustness and performance of the flight control systems to in-flight damage. A detachable wing with release mechanism is designed and manufactured to simulate 20% wing loss. The flight control system is implemented on a practical UAV and a successful flight test shows that it performs fully autonomous flight control, and is able to accommodate an in-flight partial wing loss.<br>AFRIKAANSE OPSOMMING: In hierdie tesis word die ontwerp, analise, implementasie en verifikasie van ’n fout-verdraende onbemande vliegtuig beheerstelsel wat robuust is tot strukturele skade wat die natuurlike vlug dinamika van die voertuig asimmetries maak, voorgestel. Die hoofdoel van hierdie robuuste beheer argitektuur is om stabiliteit te verseker na die skade aangerig is. Die beheerstelsel moet die skielike verandering van normale na beskadigde vlug hanteer sonder enige eksplisiete kennis daarvan. Dus word ’n robuuste beheer aanslag verkies bo ’n aanpassende beheer struktuur. Tweedens moet die vlugbeheerstelsel robuust genoeg wees om steeds die gewenste reaksietyd en aanvaarbare oorgangsverskynsels te kan hanteer, tydens beide normale en beskadigde vlug. ’n Asimmetriese ses grade van vryheid beweginsvergelykings model word afgelei. Die model het die vermoë om veranderinge in die aerodinamiese model van die vliegtuig, sowel as massamiddelpunt verskuiwing, voor te stel. “Vortex Lattice” metodes is gebruik om die aerodinamiese koëffisiënte van die beskadigde vlerk voor te stel tussen 0% en 40% verlies. ’n Sekwensiële kwadratiese programmering optimiserings algorithme is aangewend op die krag en moment vergelykings om die ekwilibrium vlug toestand en aktueerder defleksies te vind vir ’n asimmetriese vliegtuig met konstante lugspoed en hoogte. Die ekwilibrium vlug toestand word verder beperk deur ’n nul rolhoek, ’n nul sygliphoek of ’n relatiewe weging van die twee. Die bepaalde ekwilibrium defleksies word dan vergelyk met die fisiese limiete om hulle geldigheid te bepaal vir ekwilibrium vlug. As ’n geldige ekwilibrium toestand bestaan, kan die relatiewe stabiliteit van die vliegtuig se natuurlike modusse ontleed word as ’n persentasie van vlerkverlies deur die wortellokusse van die ooplus pole na te gaan. ’n Versnellings-gebaseerde vlug beheerstelsel argitektuur is ontwerp en geïmplementeer. Daarna is die robuustheid ontleed as ’n funksie van die persentasie vlerkverlies. Die robuustheid en gedrag van hierdie vlugbeheerstelsel is geverifieer met ’n nie-linêre simulasie vir 0 tot 40% vlerkverlies. Praktiese vlugtoetse is onderneem om die robuustheid en gedrag tydens/na skade gedurende ’n vlug, te verifeer. ’n Vlerkverlies meganisme is ontwerp en vervaardig om 20% vlerkverlies te simuleer. Die vlugbeheerstelsel is geïmplementeer op ’n onbemande vliegtuig en die daaropvolgende suksesvolle vlug lewer bewys dat die vlugbeheerstelsel wel skade, in die vorm van gedeeltelike vlerkverlies, tydens vlug kan hanteer.

The study of the flow of a fan blade was conducted to improve tonal fan noise reduction by optimizing an existing flow control configuration. The current configuration consisted of a trailing edge Slot with a flow control area of 0.045 in² per inch span with an exit angle of -3.3° with respect to the blade exit angle. Two other flow control configurations containing discrete jets were investigated. For the first configuration, the trailing edge jets (TEJ), the fan blade was modified with discrete jets spaced 0.3 inches apart with a flow control area of 0.01 in² per inch span positioned on the trailing edge aimed at -3.3° with respect to the blade exit angle. Similarly, discrete jets were also placed on the suction surface at 95.5% chord aimed at 15° with respect to the local blade surface. This configuration is referred to as the suction surface jet (SSJ). The discrete jets for both configurations were designed to be choked while injecting a mass flow rate of 1.00% of the fan through-flow. Computational Fluid Dynamics (CFD) was used to model new configurations and study subsequent changes in total pressure deficit using a blade design inlet Mach number of 0.73, Reynolds number based on chord length of 1.67 à 106, and design incidence angle of 0°. Experimental testing was later conducted in a 2D cascade tunnel. The TEJ and SSJ were tested at design blowing of 1.00% and at off-design conditions of 0.50%, 0.75%, and 1.25% fan through-flow. Results between the different flow control configurations were compared using a blowing coefficient. CFD showed the TEJ and SSJ offered aerodynamic improvement over the Slot configuration. Testing showed the SSJ outperformed the TEJ, as validated in CFD, producing wider and shallower wakes. SSJ area-averaged pressure losses were 25% less than TEJ at design. Noise predictions based on CFD findings showed that both TEJ and SSJ provided additional tonal sound power level attenuation over the Slot configuration at similar blowing coefficients, with the SSJ providing the most attenuation. Noise prediction based on experimental results concurred that the SSJ provided more total attenuation than the TEJ. Experimental results showed that the SSJ performed better aerodynamically and, based on analytical prediction, provided 2 dB more total attenuation than the TEJ.<br>Master of Science

Ce mémoire décrit la conception, la modélisation et la commande d’un manipulateur pneumatique,fondé sur un principe de traction aérodynamique original. De puissants jets d’air verticauxpermettent de créer un flux d’air horizontal pour manipuler des objets sans contact. Les objetssont maintenus en constante lévitation sur la surface grâce à un coussin d’air, et peuvent positionnés selon trois degrés de libert´ du plan, grâce à la combinaison adéquate et distribuéede jets d’air verticaux. Nous détaillons la conception d’un prototype original de manipulateurexploitant ce principe fluidique. Ce prototype a été intégré dans une plate-forme expérimentaleafin de valider le principe de manipulation : le système permet de déplacer des objets à unevitesse atteignant 180 mm/s. Nous avons modélisé le fonctionnement de la surface selon plusieursméthodes. Un premier modèle comportemental, fond´e sur des données expérimentales, aété établi. Il permet de simuler l’´evolution de la position d’un objet sur la surface, selon un degréde liberté . Deux modèles de connaissance, fond´es sur une étude aérodynamique fondamentale,donnent l’´evolution de la position de l’objet selon respectivement deux et trois degrés de libertédu plan. Chacun des modèles a été validé expérimentalement. Nous avons synthétisé différentscontrˆoleurs afin d’asservir la position de l’objet : un premier, de type PID, et un second, de typerobuste (méthode H1). La commande de un, puis deux degrés de liberté du système, a permisd’atteindre de bonnes performances : temps de réponse d’environ 2 s et dépassement souventinférieur à 5%. Nous avons également étudié un micro-manipulateur pneumatique permettant ded´eplacer des objets de taille millimétrique selon deux directions, grâce à des jets d’air inclinés.Ces objets peuvent atteindre des vitesses de 123 mm/s. La résolution du positionnement estinférieure à 0.4 μm<br>This thesis presents the design, the modeling and the control of a pneumatic manipulatorbased on an original aerodynamic traction principle. An horizontal air flow is induced by strongvertical air jets in order to manipulate objects without contact. The objects are maintained inconstant levitation on an air cushion. Three degrees of freedom positioning of the objects canbe realized thanks to the right combination of distributed air jets. The design of an originalmanipulator using this aerodynamic principle is detailed. The device has been integrated in anexperimental setup in order to validate the manipulation principle : objects can reach velocityof 180 mm/s. Several models of the system have been established. A first model, based on experimentaldata, gives the evolution of the 1 DOF-position of an object on the device. Twoother models, based on a fundamental aerodynamic study, respectively give the evolution of the2- and 3-DOF position of the objet. The three models have been validated experimentally. Inorder to control the position of the object, different controllers have been designed : a PID oneand a robust H1 one. The control of one and two degrees of freedom of the device gives goodperformances : settling time of around 2 s and overshoot less than 5% in most of the cases. Wehave also studied a micro-manipulator that is able to position millimetric sized objects, in twodirections, thanks to inclined air jets. Objects can reach velocity of 123 mm/s, and the resolutionof the positioning is less than 0.4 μm

Heavy industries, such as mining, generate dust in quantities that present an occupational health hazard. Prolonged exposure to the respirable dust has been found to result in many irreversible occupational ailments in thousands of miners. In underground mining applications, a variety of scrubbing systems are used to remove dust near the zones of generation. However, the wire-mesh type fibrous screens in the flooded-bed dust scrubbers used on continuous miners, are prone to clogging due to the accumulation of dust particles. This clogging results in a reduced capture efficiency and a higher exposure to the personnel. This research establishes the Vortecone, an inertial wet scrubber system, as a suitable alternative to the existing filters. The Vortecone accelerates its inlet fluids into a rapid circulatory motion into a vortex chamber, preferentially moving the heavier particles towards the impermeable surface to be trapped by the circulating water film. Vortecones are used on automobile painting lines and capture over-sprayed paint particles with cleaning efficacies exceeding 99 % while requiring only infrequent maintenance. The existing design of the Vortecone could also be altered to control the flow patterns. This dissertation presents detailed computational fluid dynamics (CFD) models to describe air flow patterns in the Vortecone in steady and transient states. Multi-phase spray models were generated to simulate injection of water into the Vortecone. The volume of fraction (VOF) approach was adopted to mimic the air-water interface. The Lagrangian particle tracking method was used to model particle capture on the interface described by the VOF. The CFD models indicate excellent cleaning efficacies, especially of larger particles. Laboratory experiments with optical measurements of aerosols in a reduced scale model of the Vortecone validate the computer models. These experiments which were performed on dust samples with particle sizes 0.3 μm and above, show that the Vortecone captures 90 % particles by mass exceeding about 5.20 and 3.20 μm at air flows of 0.28 m3/s (600 cfm) and 0.38 m3/s (800 cfm), respectively. The development of detailed large eddy simulations (LES) of air flow in the Vortecone provides a novel contribution to research by better resolving the flow patterns. An impactor-type, self-cleaning, non-clogging impingement screen system was designed as a substitute for conventional screens used in continuous miners. The screen could further be used as an efficient dust capturing mechanism with a demister in general mining applications. CFD models and laboratory experiments are presented to establish the cleaning efficacies of the system. Laboratory experiments to investigate the cleaning efficiency of a fibrous-type conventional screen is also discussed. The parameter, filter selection factor, is proposed to compare the performance of the three systems (Vortecone, fibrous screen, and impingement screen) under similar flows. The Vortecone has been found to be the most efficient dust-cleansing system, although it is the most power intensive fillter. The impingement screen shows a similar cleaning efficiency and a much higher availability compared to the conventional fibrous screen. Because of its minimal maintenance requirement, the impingement screen shows significant promise in dust-control applications in mining.

A direct-current, non-equilibrium surface glow discharge plasma in the presence of a Mach 2.85 flow is studied experimentally for flow control applications. The discharge is generated with pin-like electrodes flush mounted on a ceramic plate with sustaining currents from 25 mA to 300 mA. In the presence of a supersonic flow, two distinct discharge modes - diffuse and constricted - are observed depending on the flow and discharge operating conditions. In cathode upstream location, both diffuse and constricted discharges are observed while in cathode downstream location, the discharge mostly exhibits either constricted mode or bistable mixed mode. The effect of the discharge on the flow ("plasma actuation") is characterized by the appearance of a weak shock wave in the vicinity of the discharge. The shock is observed at low powers (~10 W) for the diffuse discharge mode but is absent for the higher power (~100 W) constricted mode. High speed laser schlieren imaging suggests that the diffuse mode plasma actuation is rapid as it occurs on a time scale that is less than 100 [mu]sec. Rotational (gas) and vibrational temperatures within the discharge are estimated by emission spectral line fits of N₂ and N⁺₂ rovibronic bands near 365-395 nm. The electronic temperatures are estimated by using the Boltzmann plot method for Fe(I) atomic lines. Rotational temperatures are found to be high (~1500 K) in the absence of a flow but drop sharply (~500 K) in the presence of a supersonic flow for both the diffuse and constricted discharge modes. The vibrational and electronic temperatures are measured to be about 3000 K and 1.25 eV (14500 K), respectively, and these temperatures are the same with and without flow. The gas (rotational) temperature spatial profiles above the cathode surface are found to be similar for the diffuse and constricted modes indicating that dilatational effects due to gas heating are similar. However, complete absence of flow actuation for the constricted mode suggests that electrostatic forces may also play an important role in supersonic plasma-flow actuation phenomena. Analytical estimates using cathode sheath theory indicates that ion pressure within the cathode sheath can be significant resulting in gas compression in the sheath and a corresponding expansion above it. The expansion in turn may fully negate the dilatational effect in the constricted case resulting in an apparent absence of forcing in the constricted case. Plasma-induced flow velocity reaches about 1 m/s in stagnant air at the discharge current of order tens of milliamps. This electrostatic forcing in the direction from anode to cathode can play an important role in the boundary layer of supersonic flow.<br>text

Moving Surface Boundary-layer Control (MSBC) was applied to several two dimensional bluff bodies using a high speed rotating cylinder as a momentum injecting device. Flow past a symmetric airfoil; a D-section; as well as square and rectangular prisms, representing a family of shapes with progressively increasing bluffness were studied in presence of the MSBC. In the case of the airfoil, the leading edge was replaced by a rotating cylinder; while the cylindrical element formed the top and bottom upstream corners of the D-section, square and rectangular prisms. Extensive wind tunnel investigation gave data about the effect of system parameters like rate of the momentum injection, angle of attack and the surface condition of the cylinder on steady and fluctuating components of the pressure distribution around the body, vortex shedding frequency (Strouhal number), and the lift and drag coefficients. A gain in the Strouhal number with increasing momentum injection suggest a decrease in the effective bluffness of the body. A significant reduction in the drag (up to 80%) was observed for the prisms at a maximum rate of momentum injection, Uc/U = 4 (Uc = cylinder surface speed, U = freestream wind speed). In the case of the airfoil, the lift coefficient increased by 160% and the stall angle was delayed from 110° to 48°. A rough criterion in terms of the location of the stagnation point was established to help decide the reversal in the direction of momentum injection as a function of angle of attack to ensure continued benefit. Effect of momentum injection in suppressing the vortex resonance and galloping type of instabilities were studied by mounting the bluff prism models on a dynamic-test rig inside the wind tunnel test section. The measurement of amplitude and frequency of the transverse oscillations over a range of wind speeds showed complete vibration suppression for momentum injection rates Uc/U < 2. Asymmetric momentum injection (e.g. top cylinder rotating, bottom cylinder stationary), was also found to be effective in disrupting the vortex shedding process and thereby inhibiting vibrations. The suppression of galloping instability in presence of the MSBC was also predicted by the quasisteady analysis. A numerical panel method was developed to simulate bluff body fluid dynamnics in presence of the MSBC. The body is descretized into a large number of panels (100 - 150) with each panel comprising of a continuous distribution of linearly varying vorticity and a constant source strength. A set of linear algebraic equations approximates the Fredholm type integral equation derived from ideal fluid flow assumption. The wake is modelled by upper and lower 'free vortex layers' emanating from the separation points on the body. Vorticity is allowed to be shed and dissipated as it is convected downstream along the panels on the 'free vortex layers'. An analytical expression relates the point vortex modelling a.rotating cylinder to the rate of momentum injection. The panel method is capable of treating multielement configurations (e.g. a rotating cylinder and the truncated airfoil). An iterative scheme based on the convergence of the wake shape is used to obtain the final solution. The numerically obtained pressure distribution, the lift and drag coefficients agree well with the experimental results. Flow visualization studies in a water channel were performed to obtain better physical insight into the MSBC process. Plexiglas models with rotating cylinders in conjunction with a fine suspension of polyvinyl chloride particles and slit lighting were used to visualize the streaklines. The still photographs and video movie recorded, rather dramatically, the effectiveness of the MSBC in suppressing separation and vortex shedding, making the flow approach the potential character. Overall, the present research firmly establishes potential of the MSBC as a versatile tool for lift augmentation, drag reduction and vibration suppression of several bluff bodies encountered in industrial engineering practice.

An aluminum NACA 0018 airfoil testbed was constructed with 95 static pressure taps and 25 embedded microphones to enable novel time-resolved measurements of surface pressure. The main objective of this investigation is to utilize time-resolved surface pressure measurements to estimate salient flow characteristics in the separated flow region over the upper surface of an airfoil. The flow development over the airfoil was examined using hot wire anemometry and mean surface pressure for a range of Reynolds numbers from 80x103 to 200x103 and angles of attack from 0° to 18°. For these parameters, laminar boundary layer separation takes place on the upper surface and two flow regimes occur: (i) separation is followed by flow reattachment, so that a separation bubble forms and (ii) separation occurs without subsequent reattachment. Measurements of velocity and mean surface pressure were used to characterize the separated flow region and its effect on airfoil performance using the lift coefficient. In addition, the transition process and the evolution of disturbances were examined. The lift curve characteristics were found to be linked to the rate of change of the separation, transition, and reattachment locations with the angle of attack. For both flow regimes, transition was observed in the separated shear layer. Specifically, the amplification of disturbances within a band of frequencies in the separated shear layer resulted in laminar to turbulent transition. Validation of time-resolved surface pressure measurements was performed for Rec = 100x103 at α = 8° and α = 12°, corresponding to regimes of flow separation with and without reattachment, respectively. A comparative analysis of simultaneous velocity and time-resolved surface pressure measurements showed that the characteristics and development of velocity fluctuations associated with disturbances in the separated shear layer can be extracted from time-resolved surface pressure measurements. Specifically, within the separated flow region, the amplitude of periodic oscillations in the surface pressure signal associated with disturbances in the separated shear layer grew in the streamwise direction. In addition, the frequency at the spectral peak of the amplified disturbances in the separated shear layer was identified. Based on the results of the validation analysis, time-resolved surface pressure measurement analysis techniques were applied for a Reynolds number range from 60x103 to 130x103 and angles of attack from 6° to 16°. Within the separated flow region, the streamwise growth of surface pressure fluctuations is distinctly different depending on the flow regime. Specifically, within the separation bubble, the RMS surface pressure fluctuations increase in the streamwise direction and reach a peak just upstream of the reattachment location. The observed trend is in agreement with that observed for other separating-reattaching flows on geometries such as the forward and backward facing step and splitter plate with fence. In contrast to the separation bubble formation, when the separated shear layer fails to reattach to the airfoil surface, RMS surface pressure fluctuations increase in the streamwise direction with no maximum and the amplitude is significantly lower than those observed in the separation bubble. Surface pressure signals were further examined to identify the frequency, convective velocity, and spanwise uniformity of disturbances in the separated shear layer. Specifically, for both flow regimes, the fundamental frequency and corresponding Strouhal number exhibit a power-law dependency on the Reynolds number. Based on the available data for which velocity measurements were obtained in the separated flow region, the convective velocity matched the mean velocity at the wall-normal distance corresponding to the maximum turbulence intensity. A distinct increase in the convective velocity of disturbances in the separated shear layer was found when the airfoil was stalled in comparison to that found in the separation bubble. From statistical analysis of surface pressure signals in the spanwise direction, it was found that disturbances are strongly two-dimensional in the laminar portion of the separated shear layer and become three-dimensional through the transition process.