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.

<p>Methods for structural design, control, and testing offlexible aircraft structures are considered. Focus is onnonconventional aircraft con- figurations and control concepts.The interaction between analysis and testing is a central topicand all studies include validation testing and comparisonbetween computational and experimental results.</p><p>The first part of the thesis is concerned with the designand testing of an aeroelastic wind-tunnel model representing aBlended Wing Body (BWB) aircraft. The investigations show thata somewhat simplified wind-tunnel model design concept isuseful and efficient for the type of investigations considered.Also, the studies indicate that well established numericaltools are capable of predicting the aeroelastic behavior of theBWB aircraft with reasonable accuracy. Accurate prediction ofthe control surface aerodynamics is however found to bedifficult.</p><p>A new aerodynamic boundary element method for aeroelastictimedomain simulations and its experimental validation arepresented. The properties of the method are compared totraditional methods as well as to experimental results. Thestudy indicates that the method is capable of efficient andaccurate aeroelastic simulations.</p><p>Next, a method for tailoring a structure with respect to itsaeroelastic behavior is presented. The method is based onnumerical optimization techniques and developed for efficientdesign of aeroelastic wind-tunnel models with prescribed staticand dynamic aeroelastic properties. Experimental validationshows that the design method is useful in practice and that itprovides a more efficient handling of the dynamic aeroelasticproperties compared to previous methods.</p><p>Finally, the use of multiple control surfaces andaeroelastic effects for efficient roll maneuvering isconsidered. The idea is to design a controller that takesadvantage of the elasticity of the structure for performancebenefits. By use of optimization methods in combination with afairly simple control system, good maneuvering performance isobtained with minimal control effort. Validation testing usinga flexible wind-tunnel model and a real-time control systemshows that the control strategy is successful in practice.Keywords: aeroelasticity, active aeroelastic structures,aeroelastic tailoring, control, structural optimization,wind-tunnel testing.</p>

A key remaining challenge in the design of hypersonic vehicles is the incomplete understanding of the process of boundary-layer transition. Turbulent heating rates are substantially higher than those for a laminar boundary layer, and large uncertainties in transition prediction therefore demand conservative, inefficient designs for thermal protection systems. It is only through close collaboration between theory, experiment, and computation that the state of the art can be advanced, but experiments relevant to flight require ground-test facilities with very low disturbance levels. To enable this work, a unique Mach 6 low-disturbance wind tunnel, previously of NASA Langley Research Center, is established within a new pressure-vacuum blow-down infrastructure at Texas A&M. A 40-second run time at constant conditions enables detailed measurements for comparison with computation. The freestream environment is extensively characterized, with a large region of low-disturbance flow found to be reliably present for unit Reynolds numbers Re < 11×10^6 m-1. Experiments are performed on a 5º half-angle flared cone model at Re = 10×10^6 m-1 and zero angle of attack. For the study of the second-mode instability, well-resolved boundary-layer profiles of mean and fluctuating mass flux are acquired at several axial locations using hot-wire probes with a bandwidth of 330 kHz. The second mode instability is observed to undergo significant growth between 250 and 310 kHz. Mode shapes of the disturbance agree well with those predicted from linear parabolized stability equation (LPSE) computations. A 17% (40 kHz) disagreement is observed in the frequency for most-amplified growth between experiment and LPSE. Possible sources of the disagreement are discussed, and the effect of small misalignments of the model is quantified experimentally. A focused schlieren deflectometer with high bandwidth (1 MHz) and high signal-to-noise ratio is employed to complement the hot-wire work. The second-mode fundamental at 250 kHz is observed, as well as additional harmonic content not discernible in the hot-wire measurements at two and three times the fundamental. A bispectral analysis shows that after sufficient amplification of the second mode, several nonlinear mechanisms become significant, including ones involving the third harmonic, which have not hitherto been reported in the literature.

During wind tunnel testing of jet-lift, short take-off and vertical landing (STOVL) aircraft it is usual to simulate the jet efflux but not the intake flows. The intakes, which are commonly faired over or are unpowered, are generally tested in separate wind tunnel experiments. The forces acting on the wind tunnel model are determined by the linear addition of the forces obtained from the two separate tests. There is some doubt as to whether this is a valid approach. A systematic experimental investigation was, therefore, conducted to determine the magnitude of any jet/intake interference effects on a generic jet-lift STOVL aircraft in transitional flight, out of ground effect. Comparisons made between separate and simultaneous jet and intake testing concluded that a mutual jet/intake interference effect does exist. The existence of this interference means that the aerodynamic wing lift loss in transitional flight deduced from isolated jet and intake testing is less than the lift loss obtained from simultaneous jet and intake testing. The experimental research was supplemented by some simplified computational fluid dynamics (CFD) studies of elements of the flow-field about the aircraft using the k-e turbulence model. The numerical modelling enabled aspects of the flow-field around the aircraft to be visualised which could not easily be done using the experimental apparatus. It is a requirement of the Eng]) programme that part of this thesis must address a management topic linked to the research. In this case the management aspects of wind tunnel project work were examined. A scenario was developed which established a requirement for a large-scale, low-speed wind tunnel with a Reynolds number capability of 20 million. A study was performed on the decision-making process and investment appraisal methods used in the procurement of such a wind tunnel.

During wind tunnel testing of jet-lift, short take-off and vertical landing (STOVL) aircraft it is usual to simulate the jet efflux but not the intake flows. The intakes, which are commonly faired over or are unpowered, are generally tested in separate wind tunnel experiments. The forces acting on the wind tunnel model are determined by the linear addition of the forces obtained from the two separate tests. There is some doubt as to whether this is a valid approach. A systematic experimental investigation was, therefore, conducted to determine the magnitude of any jet/intake interference effects on a generic jet-lift STOVL aircraft in transitional flight, out of ground effect. Comparisons made between separate and simultaneous jet and intake testing concluded that a mutual jet/intake interference effect does exist. The existence of this interference means that the aerodynamic wing lift loss in transitional flight deduced from isolated jet and intake testing is less than the lift loss obtained from simultaneous jet and intake testing. The experimental research was supplemented by some simplified computational fluid dynamics (CFD) studies of elements of the flow-field about the aircraft using the k-e turbulence model. The numerical modelling enabled aspects of the flow-field around the aircraft to be visualised which could not easily be done using the experimental apparatus. It is a requirement of the Eng]) programme that part of this thesis must address a management topic linked to the research. In this case the management aspects of wind tunnel project work were examined. A scenario was developed which established a requirement for a large-scale, low-speed wind tunnel with a Reynolds number capability of 20 million. A study was performed on the decision-making process and investment appraisal methods used in the procurement of such a wind tunnel.

The circulation control system onboard Cal Poly's Advanced Model for Extreme Lift and Improved Aeroacoustics was a critical component of a highly complex wind tunnel model produced in order to fulfill the requirements of a NASA Research Announcement awarded to David Marshall of the Aerospace Engineering Department. The model was based on a next generation, 150 passenger, regional, cruise efficient, short take-off and landing concept aircraft that achieved high lift through circulation control wings and over-the-wing mounted engines. The wind tunnel model was 10-ft in span, used turbine propulsion simulators, and had a functioning circulation control system driven from tunnel supplied high pressure air. Wind tunnel test results will be compiled into an open-source database intended for validation of predictive tools whose purpose is to advance the state- of-the-art in predictive capabilities for the next generation aircraft configurations. The model's circulation control system produced highly directional, nonuniform flow, and required significant modification in order to generate flow suitable for representation in predictive software. The effort and methods used to generate uniform flow along the circulation control slots is detailed herein. Additionally the results of the system characterization are presented and include a thorough analysis of the slot height, the wing symmetry, and total pressure at the circulation control jet exit. These datasets are intended to aid in making adjustments to the simulation such that it accurately reflects the condition at which the model was tested. Many flow visualization results from the wind tunnel test are also presented to serve as a medium of comparison for results from predictive tools. Oil flow visualization was conducted at many test conditions and provides insight to AMELIA's surface flow in blown and unblown regions. Of particular interest were streamlines at the wingblend, which exhibited some outboard turning, and streamlines on the lower surface where the leading edge stagnation point was investigated. Smoke flow visualization was also utilized to explore the flowfield. The deflection of a individual streamline, under the influence of a changing discharge coefficient as investigated along with the discharge coefficients effect on the extended flowfield. Collectively, the images depict the massive augmentation of the flowfield caused by the presence of the circulation control wing.

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As technology for manufacturing small scale prototypes of aeroplanes has become cheaper and more easily viable, the process of Rapid Prototyping has become more common. Rapid Prototyping allows for the fundamental aerodynamic qualities of a geometric body to be tested in a wind tunnel using a small scale prototype. This means smaller prototypes of aircraft can be manufactured more rapidly and at a lower cost, allowing for more extensive testing of a design’s final aerodynamic qualities before any actual full-size production. In order to gain a better insight in the behaviours of the full-sized aircraft itself, a downscaled version of the KTH project UAV ALPHA has been deigned for testing in a low speed wind tunnel. The design will be used in further testing to help confirm simulations and estimations done on the ALPHA of its aerodynamic performance.

Investigation of complex multirotor aerodynamic phenomena via wind tunnel experimentation is becoming extremely important with the rapid progress in advanced distributed propulsion VTOL concepts. Much of this experimentation is being performed in large, highly advanced tunnels. However, the proliferation of this class of vehicles extends to small aircraft used by small businesses, universities, and hobbyists without ready access to this level of test facility. Therefore, there is a need to investigate whether multirotor vehicles can be adequately tested in smaller wind tunnel facilities. A test rig for a 2.82-pound quadcopter was developed to perform powered testing in the Cal Poly Aerospace Department’s Low Speed Wind Tunnel, equipped with a 3-foot tall by 4-foot wide test section. The results were compared to data from similar tests performed in the U.S. Army 7-by 10-ft Wind Tunnel at NASA Ames. The two data sets did not show close agreement in absolute terms but demonstrated similar trends. Due to measurement uncertainties, the contribution of wind tunnel interference effects to this discrepancy in measurements was not able to be properly quantified, but is likely a major contributor. Flow visualization results demonstrated that tunnel interference effects can likely be minimized by testing at high tunnel speeds with the vehicle pitched 10-degrees or more downward. Suggestions towards avoiding the pitfalls inherent to multirotor wind tunnel testing are provided. Additionally, a modified form of the conventional lift-to-drag ratio is presented as a metric of electric multirotor aerodynamic efficiency.

An experiment was performed in the Cal Poly Mechanical Engineering 2x2 ft wind tunnel to quantify the effect of spanwise synthetic jet actuation (SJA) on the drag of a NACA 0015 semispan wing. The wing, which was designed and manufactured for this experiment, has an aspect ratio of 4.20, a span of 0.427 m (16.813”), and is built around an internal array of piezoelectric actuators, which work in series to create a synthetic jet that emanates from the wingtip in the spanwise direction. Direct lift and drag measurements were taken at a Reynolds Number of 100,000 and 200,000 using a load cell/slider mechanism to quantify the effect of actuation on the lift and drag. It was found that the piezoelectric disks used in the synthetic jet actuators cause structural vibrations that have a significant effect on the aerodynamics of the NACA 0015 model. The experiment was performed in a way as to isolate the effect of vibration from the effect of the synthetic jet on the lift and drag. Lift and drag data was supported with pressure readings from 60 pressure ports distributed in rows along the span of the wing. Oil droplet flow visualization was also performed to understand the effect of SJA near the wingtip. The synthetic jet and vibration had effects on the drag. The synthetic jet with vibration decreased the drag only slightly while vibration alone could decrease drag significantly from 11.3% at α = 4° to 23.4% at α = 10° and Re = 100,000. The lift was slightly increased with a slight increase due to the jet and showed a slight increase due to vibration. Two complete rows of pressure ports at 2y/b = 37.5% and 85.1% showed changes in lift due to actuation as well. The synthetic jet increased the lift near the wingtip at 2y/b = 85.1% and had little to no effect inboard at the 37.5% location, hence, the synthetic jet changes the lift distribution on the wing. Oil flow visualization was used to support this claim. Without actuation, the footprint of the tip vortex was present on the upper surface of the wing. With actuation on, the footprint disappeared suggesting the vortex was pushed off the wingtip by the jet. It is possible that the increased lift with actuation can be caused by the vortex being pushed outboard.