An experimental investigation was performed to understand the development and suppression of the secondary flow structures within a compact, serpentine jet engine inlet duct. By employing a variety of flow diagnostic techniques, the formation of a pair of counter-rotating vortices was revealed. A modular fluidic actuator system that would apply several different methods of flow control was then designed and manufactured to improve duct performance. At the two bends of the inlet, conformal flow control devices were installed to deliver varying degrees of boundary layer suction, suction and steady fluid injection, and suction and oscillatory injection. Testing showed that suction alone could delay flow separation and improve the pressure recovery of the duct by as much as 70%. However, this technique was not able to rid the duct completely of the nonuniformities that exist at the engine face plane. Suction with steady blowing, however, increased pressure recovery by 37% and reduced distortion by 41% at the engine face. Suction with pulsed injection had the least degree of success in suppressing the secondary flow structures, with improvements in pressure recovery of only 16.5% and a detrimental impact on distortion. The potential for gains in the aerodynamic efficiency of serpentine inlets by active flow control was demonstrated in this study.

Active flow control has been extensively explored in wind tunnel studies but successful in-flight implementation of an active flow control technology still remains a challenge. This thesis presents implementation of active flow control technology onboard a 33% scale Extra 330S ARF aircraft, wind tunnel studies and flight testing of fluidic actuators. The design and construction of the pulsed blowing system for stall suppression (LE actuator) and continuous blowing system for roll control (TE actuator) and pitch control have been presented. Full scale wind tunnel testing in 7̕ X 10 Oran W. Nicks low speed wind tunnel shows that the TE actuators are about 50% effective as the conventional ailerons. The LE actuator is found to be capable of suppressing stall from 12° to about 22°. Comparison of characteristics of Active elevator and conventional elevator in 3' X 4' low speed wind tunnel show that, the active elevator is as effective as of conventional elevator deflected at 5°. Flight tests show that TE actuators are able to control the aircraft in flight in banked turns. The measured roll rates in-flight support the wind tunnel test findings.

Laminar flow control has been studied for several decades in an effort to achieve higher efficiencies for aircraft. Successful implementation of laminar flow control technology on transport aircraft could significantly reduce drag and increase operating efficiency and range. However, the crossflow instability present on swept-wing boundary layers has been a chief hurdle in the design of laminar wings. The use of spanwise-periodic discrete roughness elements (DREs) applied near the leading edge of a swept-wing typical of a transport aircraft represents a promising technique able to control crossflow and delay transition to accomplish the goal of increased laminar flow. Recently, the Flight Research Laboratory at Texas A&M University conducted an extensive flight test study using DREs on a swept-wing model at chord Reynolds numbers in the range of eight million. The results of this study indicated DREs were able to double the laminar flow on the model, pushing transition back to 60 percent chord. With the successful demonstration of DRE technology at these lower chord Reynolds numbers, the next logical step is to extend the technology to higher Reynolds numbers in the range of 15 to 20 million typical of smaller transport aircraft. To conduct the flight tests at the higher Reynolds numbers, DREs will be placed on a wing glove attached to the aircraft wing. However, a feasibility study was necessary before initiating the flight-testing. First, a suitable aircraft able to achieve the Reynolds numbers and accommodate a wing glove was identified. Next, a full CFD analysis of the aircraft was performed to determine any adverse effects on the wing flow-field from the aircraft engines. This required an accurate CAD model of the selected aircraft. Proper modeling techniques were needed to represent the effects of the aircraft engine. Once sufficient CFD results were obtained, they were used as guidance for the placement of the glove. The attainable chord Reynolds numbers based on the recommendations for the wing glove placement then determined if the selected aircraft was suitable for the flight-testing.

In this dissertation, we present performance prediction models for rate-based and window¡Vbased flow control mechanisms. For rate-based flow control, such as in ATM network, we derive two analytical models to predict the ACR rates for congestion-free and congestion networks, respectively. To coordinate the cooperative problems of TCP over ATM networks, we propose a new algorithm to monitor the states of ATM switches and adjust TCP congestion window size based on RM cells. For window-based flow control mechanisms, such as in TCP-Reno and TCP-SACK, we respectively present analytical models to systematically capture the characteristics of multiple consecutive packet losses in TCP windows. Through fast retransmission, the lost packets may or may not be recovered. Thus, we present upper bound analyses for slow start and congestion avoidance phases to study the effects of multiple packet losses on TCP performance. Above the proposed upper bounds, the lost packets may not be successfully recovered through fast retransmission. Finally, we develop a model to study the TCP performance in terms of throughput degradation resulted from multiple consecutive packet losses. The analytical results from the throughput degradation model are validated through OPNET simulation.

In this work, an active flow control method is studied numerically by using a synthetic jet over a NACA 0015 airfoil. Unsteady, turbulent flows over the NACA 0015 airfoil are computed using a Navier-Stokes solver. The Spalart-Allmaras turbulence model is employed in all computations. Unsteady flow solutions are computed in parallel using Parallel Virtual Machine library routines in a computer cluster. The synthetic jet is implemented to the flow solver as a boundary condition. Response Surface Methodology is employed for the optimization of synthetic jet parameters at various angles of attack. The synthetic jet parameters / the jet velocity, the jet location, the jet angle and the jet frequency are optimized to maximize the lift to drag ratio. The optimization study is performed for a constant value of jet power coefficient. The jet slot size is used as a dependent parameter in the optimization studies. The optimization study has shown that the jet velocity and the jet location are the dominant synthetic jet parameters. The optimum synthetic jet angle is observed to be increasing as the angle of attack increases. The optimum jet location is observed to be moving through the leading edge as angle of attack increases for the separated flows. It is observed that the application of the synthetic jet delays the flow separation on the suction side of the airfoil and increases the lift to drag ratio significantly especially at post stall angles of attack. The application of the synthetic jet is observed to be less effective for attached flows.

Active flow control by a jet over a 12.5% thick elliptic profile is investigated numerically. Unsteady flowfields are calculated with a Navier Stokes solver. The numerical method is first validated without the jet and with the presence of steady-blowing and pulsating jets. Three jet types, namely steady, pulsating and synthetic jets, are next compared with each other and it is shown that the most drag reduction is achieved by a synthetic jet and the most lift enhancement is achieved by a steady jet. The influences of the jet location, the jet velocity, the jet frequency, the jet slot length and the jet angle on the flowfield is parametrically studied. It is shown that the jet location and the jet velocity are the most effective parameters. The jet parameters are optimized to minimize the drag coefficient while keeping the jet power constant. The drag is reduced by 32.5% for the angle of attack 0 and by 24% for the angle of attack 4.

This experiment is to investigate the effects of fluid with and without mass injection through a slit on the vortex shedding from a single cylinder. We research Reynolds Numbers on ranges from 800 to 4000. We used four kinds of ways which contain no mass injection, steady blowing, steady suction and oscillatory jet to study of the wake behind a circular cylinder under excitation in this experiment. No mass injection is measured for the sake of its reliability and comparability of experiment. Steady blowing and suction are applied to influence the wake flow. An oscillatory jet is used to influence the wake flow with varying frequencies and amplitudes. The experiment looks forward to use the results of this experiment so as to research into the effects on the wake flow, including the velocity values of fluctuation and turbulence intensity of the vortices structure, the dominant frequency in the flow pattern on a single cylinder. The main parameters in the study are the frequency, momentum and the location of the blowing and suction jet, which are a steady jet or unsteady oscillatory jet. Flow visualization has been carried out to investigate the interaction of steady or unsteady fluid perturbation and the vortex shedding of a cylinder.

The objective of this research is to present a theoretical study of the compressibility effects on the performance of an electric D/C motor driven synthetic jet actuator for flow separation control. Hot wire anemometer experiments were conducted to validate the jet exit velocities predicted by the theoretical model. The optimal jet exit velocity required to achieve maximum flow reattachment at reasonable blowing momentum coefficients is predicted. A dynamic electro-acoustic model of the D/C motor driven actuator is developed to accurately predict its performance and efficiency. This model should help formulate a feedback optimal control strategy for real-time flow control using an array of actuators. This model is validated by comparing with hot wire anemometer experiments conducted under similar conditions. The effects of geometric parameters like the slot width, slot geometry, and cavity volume on the performance of the actuator are also tested using this model.

The aerodynamic steering forces and moments on a wind tunnel model of an axisymmetric bluff body are altered by induced segmented attachment of the separated flow over an azimuthal Coanda surface. The model is suspended in the wind tunnel by eight thin wires for minimal support interference within the wake. Each wire is instrumented with a miniature strain gage sensor for direct dynamic force measurements. Control is effected by an array of synthetic jet actuators that emanate from narrow, azimuthally-segmented slots, within a backward facing step. The aerodynamic effects are characterized using hot-wire anemometry and PIV measurements. In the first set of experiments, the array of synthetic jets is distributed around the perimeter of the circular tail end which is extended into a Coanda surface. The fluidic actuation results in segmented vectoring of the separated base flow along the rear Coanda surface and induces asymmetric aerodynamic forces and moments that can effect steering during flight. Transitory modulation of the actuation waveform of multiple actuators around the tail leads to the generation of significant dynamic side forces of controlled magnitude and direction with the potential utility for flight stabilization and fast maneuvering. In a second set of experiments the array of the synthetic jets is placed upstream of a mid-body axisymmetric cavity. A single jet induces a quasi-steady, nearly-matched force couple at the upstream and downstream ends of the cavity. Furthermore, transitory activation of multiple jets can be used to control the onset and sequencing of the couple forces and therefore the resultant force and moment.

The aerodynamic forces and moments on an axisymmetric body of revolution are controlled in a low-speed wind tunnel by induced local flow attachment. Control is effected by an array of aft-facing synthetic jets emanating from narrow, azimuthally segmented slots embedded within an axisymmetric backward facing step. The actuation results in a localized, segmented vectoring of the separated base flow along a rear Coanda surface and induced asymmetric aerodynamic forces and moments. The observed effects are investigated in both quasi-steady and transient states, with emphasis on parametric dependence. It is shown that the magnitude of the effected forces can be substantially increased by slight variations of the Coanda surface geometry. Force and velocity measurements are used to elucidate the mechanisms by which the synthetic jets produce asymmetric aerodynamic forces and moments, demonstrating a novel method to steer axisymmetric bodies during flight.