A long aspect ratio synthetic jet is produced through an axial slit along part of the length of a cylinder. The jet is excited acoustically by a pair of loudspeakers mounted at the cylinder terminations. The study compares between the performance of two different slits with aspect ratios of 273 and 773. The comparison is based on the spanwise distribution of the mean jet velocity and phase between the jet velocity fluctuations and the excitation signal. Three different frequencies and amplitudes are used to excite the speakers covering the range of frequencies used in the control application. For both cases studied the mean centerline velocity of the jet increases with increasing the amplitude of the exciting signal, but decreases with increasing its frequency. Moreover, velocity deficits of up to 30% are evident as the midspan of the cylinder is approached from either end. Similar trends are also observed for the centerline phase distributions of the velocity fluctuations, with deficits of up to 130°. However, it is observed that for the long slit case the deficits in both the velocity and phase distributions are much larger than those for the short one. The synthetic jet is then mounted in the upstream cylinder of a tandem cylinder arrangement to be used as a control actuator for controlling the vibrations of the downstream cylinder. A simple feedback control mechanism is used at a Reynolds number of about 6.3x104. This Reynolds number corresponds to the case' where the downstream cylinder’s response is dominated with two frequency components, one at the resonance frequency of the cylinder, which is excited by broadband turbulence in the flow, and the other at the vortex shedding frequency. Both slits studied for the characterization experiments are used to compare their performance as control actuators. Both jets produce comparable reductions in the vibration of the downstream cylinder. A reduction of about 20% in the total RMS amplitude of the vibrations signal is achieved. This amounts to a reduction of about 50% in the resonant peak and an average value of about 40% in the vortex shedding peak. The optimal values of gain and time lag of the controller are then used to investigate the effect of the jet on the flow. It is found that the short slit jet produced an effect that was traced up to 1.875 diameters downstream, while the effect of the long slit jet dropped dramatically very close to the upstream cylinder. / Thesis / Master of Applied Science (MASc)

A series of flight tests was performed using a swept-wing model mounted on a Cessna O-2 aircraft. The crossflow waves on the airfoil were excited by pneumatic spanwise-periodic distributed roughness elements (DREs). The objective of the experiment was to determine the roughness receptivity i.e. the relationship between roughness height and the amplitude of the unstable crossflow wave. The local skin-friction variation was measured using an array of calibrated and temperature-compensated hotfilm sensors. The amplitudes of the disturbance shear stress were compared to the amplitudes of the DREs. It was found that there is a relationship between the shear stress and DRE amplitude that needs to be studied more before any definitely conclusions can be made. It was also found that the sensitivity of the crossflow to DREs is highly dependent on the freestream turbulence levels.

This thesis considers the control of combustion oscillations, motivated by the susceptibility of lean premixed combustion to such oscillations, and the long and expensive development and commissioning times that this is giving rise to. The controller used is both closed-loop, employing an actuator to modify some system parameter in response to a measured signal, and adaptive, meaning that it is able to maintain control over a wide range of operating conditions. The controller is applied to combustion systems with annular geometries, where instabilities can occur both longitudinally and azimuthally, and which require multiple sensors and multiple actuators for control. One of the requirements of Lyapunov-based adaptive control which is particularly troublesome for combustion systems is then addressed: that the sign of the high-frequency gain of the open-loop system is known. We address it by using an adaptive controller which employs a Nussbaum gain, and successfully apply it experimentally to combustion oscillations in a Rijke tube. Another type of fluid-acoustic resonance is then considered: the compressible flow past a shallow cavity. We start by finding a linear model of the cavity flow's dynamics, or its 'transfer function', which we identify from direct numerical simulations. We compare this measured transfer function to that given by a conceptual model which is based on the Rossiter mechanism, and which models each component of the flow physics separately. We then look at using closed-loop control to eliminate these cavity oscillations. We start by designing a robust H₂ controller based on a balanced reduced order model of the system, the model being provided by the Eigensystem Realization Algorithm (ERA). The robust controller provides closed-loop stability over a much wider Mach number range than seen in previous studies. Finally, we look at the suitability of the adaptive controller, earlier developed for combustion oscillations, for the cavity. Based on some general properties of the cavity flow, and by using collocated control, the oscillations are eliminated at all Mach numbers tested in the range 0.4 ≤ M ≤ 0.8.

Aerodynamic flow control using internal acoustic excitation holds promise as it combines the simplicity of passive flow control techniques (in terms of added weight and operational complexity) with the control authority of active flow control methods. While previous studies have analyzed the effects of acoustic excitation on steady wing aerodynamics, the effect of excitation on the unsteady aerodynamics is not known, which is the aim of the current effort. Internally mounted speakers on a symmetric National Advisory Committee for Aeronautics (NACA) 0012 wing are used to excite the unsteady boundary layer at the wing’s leading edge as it executes linear pitch motions ranging from quasi-steady (trailing-edge driven stall) to vortex dominated (mixed leading- and trailing-edge driven stall) motions at freestream Reynolds numbers (����) of 120, 000 and 180, 000. Experimental results show that, while acoustic excitation delays stall for quasi-steady motions, it enhances lift in the linear region and increases leading-edge vortex strength for vortex -dominated motions. The degree of change was observed to be a function of the excitation frequency. The current work establishes the effects of acoustic flow excitation in unsteady, low-���� wing aerodynamics and provides insights on the path forward to effectively implement the method for active flow control.

A pulsed-plasma jet (called a "spark-jet" by other researchers), is a high-speed synthetic jet that is generated by striking an electrical discharge in a small cavity. The gas in the cavity pressurizes owing to the heating and is allowed to escape through a small orifice. A series of experiments were conducted to determine the characteristics of the pulsed-plasma jet issuing into stagnant air at a pressure of 45 Torr. These results show that typical jet exit velocities of about 250 m/s can be induced with discharge energies of about 30 mJ per jet. Furthermore, the maximum pulsing frequency was found to be about 5 kHz, because above this frequency the jet begins to misfire. The misfiring appears to be due to the finite time it takes for the cavity to be recharged with ambient air between discharge pulses. The velocity at the exit of the jet is found to be primarily dependent on the discharge current and independent of other discharge parameters such as cavity volume and orifice diameter. Temperature measurements are made using optical emission spectroscopy and reveal the presence of considerable non-equilibrium between rotational and vibrational modes. The gas heating efficiency was found to be 10% and this parameter is shown to have a direct effect on the plasma jet velocity. These results indicate that the pulsed-plasma jet creates a sufficiently strong flow perturbation that is holds great promise as a supersonic flow actuator. An experimental study is conducted to characterize the performance of a pulsed-plasma jet for potential use in supersonic flow control applications. To obtain an estimate of the relative strength of the pulsed-plasma jet, the jet is injected normally into a Mach 3 cross-flow and the penetration distance is measured by using schlieren imaging. These measurements show that the jet penetrates 1.5 [delta], where [delta] is the boundary layer thickness, into the cross-flow and the jet-to-crossflow momentum flux ratio is estimated to be 0.6. An array of pulsed-plasma jets was issued from different locations upstream of a 30-degree compression ramp in a Mach 3 flow. Furthermore, two different jet configurations were used: normal injection and pitched and skewed injection. The pitched and skewed configuration was used to see if the jets could act as high-bandwidth pulsed vortex generators. The interaction between the jets and the separation shock was studied using phase-locked schlieren imaging. Results show that the plasma jets cause a significant disturbance to the separation shock and clearly influence its unsteadiness. While all plasma jet configurations tested caused an upstream motion of the separation shock, pitched and skewed plasma jets caused an initial downstream shock motion before the upstream motion, demonstrating the potential use of these plasma jets as vortex generator jets. The effect of the plasma jet array on the separation shock unsteadiness is studied in a time-resolved manner by using 10 kHz schlieren imaging and fast-response wall pressure measurements. An array of three pulsed-plasma jets, in a pitched and skewed configuration, is used to force the unsteady motion of the interaction formed by a 24° compression ramp in a Mach 3 flow. The Reynolds number of the incoming boundary layer is Re[theta]=3300. Results show that when the pulsed jet array is placed upstream of the interaction, the jets cause the separation shock to move in a quasi-periodic manner, i.e., nearly in sync with the pulsing cycle. As the jet fluid convects across the separation shock, the shock responds by moving upstream, which is primarily due to the presence of hot gas and hence the lower effective Mach number of the incoming flow. Once the hot gases pass through the interaction, the separation shock recovers by moving downstream, and this recovery velocity is approximately 1% to 3% of the free stream velocity. With forcing, the low-frequency energy content of the pressure fluctuations at a given location under the intermittent region decreases significantly. This is believed to be a result of an increase in the mean scale of the interaction under forced conditions. Pulsed-jet injection are also employed within the separation bubble, but negligible changes to the separation shock motion were observed. These results indicate that influencing the dynamics of this compression ramp interaction is much more effective by placing the actuator in the upstream boundary layer. / text

Active control of the flow over an airfoil is an area of heightened interest in the aerospace community. Previous research on flow control design processes heavily depended on trial and error and the designers knowledge and intuition. Such an approach cannot always meet the growing demands of higher design quality in less time. Successful application of computational fluid dynamics (CFD) to this kind of control problem critically depends on an efficient searching algorithm for design optimization. CFD in conjunction with Genetic Algorithms (GA) potentially offers an efficient and robust optimization method and is a promising solution for current flow control designs. But the traditional binary GA and its operators need to be transformed or re-defined to meet the requirements of real world engineering problems. Current research has combined different existing GA techniques and proposed a realcoded Explicit Adaptive Range Normal Distribution (EARND) genetic algorithm with diversity control to solve the convergence problems. First, a traditional binary-coded GA is replaced by a real-coded algorithm in which the corresponding design variables are encoded into a vector of real numbers that is conceptually closest to the real design space. Second, to address the convergence speed problem, an additional normal distribution scheme is added into the basic GA in order to monitor the global optimization process; meanwhile, design parameters boundaries are explicitly updated to eliminate unnecessary evaluations (computation) in un-promising areas to balance the workload between the global and local searching process. Third, during the initial 20% evolution (search process), the diversity of the individuals within each generation are controlled by a formula in order to conquer the problem of preliminary convergence to the local optimum. In order to better understand the two-jet control optimization results and process, at first, a single jet with a width of 2.5% the chord length is placed on a NACA 0012 airfoils upper surface simulating the blowing and suction control under Re=500,000 and angle of attack 18 degree. Nearly 300 numerical simulations are conducted over a range of parameters (jet location, amplitude and angle). The physical mechanisms that govern suction and blowing flow control are determined and analyzed, and the critical values of suction and blowing locations, amplitudes, and angles are discussed. Moreover, based on the results of single suction/blowing jet control on a NACA 0012 airfoil, the design parameters of a two-jet system are proposed. Our proposed algorithm is built on top of the CFD code, guiding the movement of two jets along the airfoils upper surface. The reasonable optimum control values are determined within the control parameter range. The current study of Genetic Algorithms on airfoil flow control has been demonstrated to be a successful optimization application.

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.

There is currently growing interest in the field of Micro Air Vehicles (MAVs). A MAV is characterized by its low Reynolds numbers flight regime which makes lift and thrust creation a significant challenge. One possible solution inspired by nature is flapping flight, but instead of the large-amplitude low-frequency motion suited to the muscular actuators of nature, small-amplitude high-frequency motion may be more suitable for electrical actuators. In this thesis the effect of small-amplitude high-frequency motion is experimentally investigated focusing on three aspects: general performance improvement, deflected jets, and the effect of geometryResults presented herein demonstrate that using small-amplitude high-frequency plunging motion on a NACA 0012 airfoil at a post-stall angle of attack of 15° can lead to significant thrust production accompanying a 305% increase in lift coefficient. At low Strouhal numbers vortices form at the leading-edge during the downward motion and then convect into the wake. This ‘mode 1’ flow field is associated with high lift but low thrust. The maximum lift enhancement was due to resonance with the natural shedding frequency, its harmonics and subharmonics. At higher Strouhal numbers the vortex remains over the leading-edge area for a larger portion of the cycle and therefore loses its coherency through impingement with the upward moving airfoil. This ‘mode 2’ flowfield is associated with low lift and high thrust. At angles of attack below 12.5° very large force bifurcations are observed. These are associated with the formation of upwards or downwards deflected jets with the direction determined by initial conditions. The upwards deflected jet is associated with the counter-clockwise Trailing Edge Vortex (TEV) loitering over the airfoil and thereby pairing with the clockwise TEV to form a dipole that convects upwards. It therefore draws fluid from the upper surface enhancing the upper surface vortex leading to high lift. The downwards deflected jet is associated with the inverse. Deflected jets were not observed at larger angles of attack as the asymmetry in the strength of the TEVs was too great; nor at smaller amplitudes as the TEV strength was insufficient. To understand the effect of geometry comparable experiments were performed for a flat plate geometry. At zero degrees angle of attack deflected jets would form, as for the NACA 0012 airfoil, however their direction would switch sinusoidally with a period on the order of 100 cycles. The lift coefficient therefore also switched. At 15° angle of attack for Strouhal numbers up to unity the performance of the flat plate was comparable to the NACA 0012 airfoil. Above unity, the upper surface and lower surface leading-edge vortices form a dipole which convects away from the upper surface resulting in increased time-averaged separation and reduced lift.

OFS provisions bandwidth in the granularity of one wavelength. With such a coarse granularity, most applications including video download, HDTV, 3D movie, and 3D TV etc. will have very short flow sizes, in the order of seconds or even sub-second, which brings challenges to the utilization efficiency of bandwidth capacity. In this thesis we study the performance of OFS for short flows. The constraint of network resources is investigated. The effect of destination and path blocking is studied. A distributed scheduling protocol called Multi-Destination Control Protocol (MDCP) is proposed to deal with such constraint. Both single wavelength and multi-wavelength configurations are studied and characterized. Simulation results demonstrate that MDCPcan improve the OFS network throughput significantly and can be as much as eighty to one hundred percent for a single-wavelength OFS network. Even for an OFS network with four wavelengths, the throughput improvement can still approach 40%. / The Internet traffic has been growing tremendously. China Telecom predicts that the compound annual growth rate of IP traffic for the next decade is at 56% - 80% and the backbone capacity will grow by another two orders of magnitudes. Furthermore, the power consumption incurred by the next generation of huge electronic IP packet switching routers in the backbone will exceed gigawatts. In view of the grave enviromnental concerns, there is a great need for a more efficient way of transporting and switching the bits. This thesis investigates a new all-optical networking technology called optical flow switching (OFS). OFS bypasses electronic routers, and provides end-to-end transparent connections, thus taking full advantage of the enormous transmission capacity of optical networks and enjoying the extremely low error rate of transparent data transmission. The most important point about OFS is that it reduces the electrical power consumption by off-loading the huge electronic routers, which could be a major constraint for future Internet growth. Unlike many other exotic all-optical switching technologies, OFS is immediately deployable using the current optical technologies, Therefore OFS is very attractive for the next generation optical networks. / Qian, Zhengfeng. / Adviser: Kwok-wai Cheung. / Source: Dissertation Abstracts International, Volume: 73-04, Section: B, page: . / Thesis (Ph.D.)--Chinese University of Hong Kong, 2011. / Includes bibliographical references (leaves 113-118). / Electronic reproduction. Hong Kong : Chinese University of Hong Kong, [2012] System requirements: Adobe Acrobat Reader. Available via World Wide Web. / Electronic reproduction. [Ann Arbor, MI] : ProQuest Information and Learning, [201-] System requirements: Adobe Acrobat Reader. Available via World Wide Web. / Abstract also in Chinese.

The Dielectric Barrier Discharge (DBD) plasma actuators have received a significant attention of many researchers in the last few decades. The main focus of these studies has been on the flow control areas such as turbulent boundary layer separation and turbulent skin friction reduction. Little attention has been paid on the effect of the DBD plasma actuators on the aerodynamic noise reduction. In this regard, the aim of the current work is to investigate the effect of the DBD plasma actuator driven at relatively low voltages on vortex-induced noise. The first part of the current work includes an extensive assessment of the effect of the DBD plasma actuator on the narrowband tonal noise radiated from a flat plate with blunt trailing edge and an airfoil (NACA 0012) with blunt and cut-in type serrated trailing edge. The measurements were carried out at Reynolds numbers between 0.75 x 10 to the power of 5 and 4 x 10 to the power of 5. It is found that the DBD plasma actuator effectiveness depends on the direction of the generated electric wind. For example, a high reduction in the narrowband tonal noise level is achieved when a direct streamwise electric wind is injected into the wake region. However, using a plasma actuator, which can induce streamwise vortices into the wake region, shows more superior noise reduction capability at lower voltages. Flow measurement results revealed that the mechanism responsible for the narrowband tonal noise reduction when the electric wind is directly injected into the wake is not due the momentum injection into the wake deficit. Rather, the streamwise jet isolates the two separated shear layers and prevents the interaction between them. On the other hand, it is found that the break-up of the spanwise coherence of the vortex shedding is responsible for the significant reduction in the tonal noise level when the spanwise actuation is used. The second part of the current work comprises the effect of the DBD plasma actuator on both the narrowband tonal noise and interaction broadband noise radiated from both single and tandem cylinder, respectively. The experiments were conducted at subcritical Reynolds number ReD = 1.1 x 10 to the power of 4. The actuators were positioned at different azimuthal angles 27° ≤ θj ≤ 153°. For the single cylinder case, the acoustic results show the DBD plasma actuator that is positioned at θj = 133° leads to a more reduction in the narrowband tonal noise level when compared to the other angles. It is found that the streamwise jet produced by the plasma actuators plays an important role in prevention of the interaction between the shear layers that separates from the cylinder. For the tandem cylinders case, the acoustic results show that the simultaneous actuation of both the upstream and the downstream cylinders leads to more reduction in both the narrowband tonal noise and the interaction broadband noise level compared with the case where only the upstream or the downstream cylinder is actuated. The mechanism responsible for this noise reduction is found to be mainly due to the streamwise jet induced by the upstream cylinder activation, which delays the vortex shedding formation and reduces the turbulence intensity in the near wake region. On the other hand, the plasma induced jet against the main-flow direction works as a virtual fluidic barrier which displaces the wake produced by the upstream cylinder away from the downstream cylinder.