Experimental Aerodynamics

For the course of aerodynamic it's planned to do some experimental experience with a small wind tunnel. In this case, the target is the profile of pressure around an infinite NACA 23012 foil  that interact with a NACA 4412 foil on which is measured Lift and Drag with a scale.
The experiment will underline the feature of a very low subsonic field, where the aft can interact and change the condition of the fwd item.
The paper is in italian, because its' my original paper, however it's not complicate and all steps are clearly define.

For the course of aerodynamic it's planned to do some experimental experience with a small wind tunnel. In this case, the target is the profile of pressure around an infinite cylinder.
The paper is in italian, because its' my original paper, however it's not complicate and all steps are clearly define.

In this thesis, the theoretical and the experimental approaches were performed to study the transitional characteristics of aerodynamic coefficients and flow field of the NACA
633-018 airfoil at Re ranging from 1.32×10E4 ~ 1.09×10E5. For theoretical approach, the catastrophe theory was applied to model the aerodynamic transition problem and to predict
some aerodynamic characteristics undiscovered before. The so-called catastrophe theory describes the discontinuous phenomena with continuous functions based on the study of
singularities of nonlinear functions. Results of aerodynamic loading measurement well verified the validity of the catastrophe theory on the aerodynamically transitional problems and flow visualization provided some physical essences of the flow field about the theory employed here. Finally, the aerodynamic transition of the airfoil was redefined and reexamined through the cusp catastrophe model.
Combining the mathematical model and the experimental results, it can be deduced that the transition and sudden change of aerodynamic coefficients, in essence, is a
phase-changing process of the topological structure of the flow field. The predicted sudden-changing points of the aerodynamic coefficients match very well with the experimental results within the error margin of α less than 0.5 degree and Re less than two thousand. It indicates that the lift or drag and Re as well as α actually constitutes a cubic hypersurface relationship during the transitional process, which is believed to be a new direction of related aerodynamic researches.

Nowadays, there are various ways to study gas and fluid flows in wind tunnels and are still an important field of scientific research. The objectives of the research intended to present the basic principle of flow visualization techniques in wind tunnels and to simulate, visualize, observe, and measure how the flow around the object.

This paper is an attempt to summarize the effect of wing tip devices employed by birds, as well as aeronautical engineers in the past to improve the performance characteristics of aircraft. The focus is on reduction of the induced drag or drag due to lift also known as inviscid drag. This paper will provide an insight on both biomimetic and conventional wing tip approaches to reduce the induced drag. Prior analysis and experiments on the aerodynamics of airplane performance due to both these separate studies have been discussed. The needs of the industry and their past inventions have been described briefly as well as new areas to be explored have been highlighted.

Particle image velocimetry was used in a low-speed wind tunnel to investigate the vortex structures of a slender reverse delta wing at various angles of attack and roll. This work investigates the characteristics of the vortices generated downstream in planes perpendicular to the freestream direction and their dependence on angles of attack and roll at a chord-based Reynolds number of Re c 3:82 10 5 . The peak tangential velocities at > 5 show a trend similar to a delta wing. A six-component balance was used to obtain the aerodynamic coefficients for a reverse delta wing as well as a delta wing for comparison. A simulation of the streamlines, velocity vectors, and surface pressure contours was carried out using computational fluid dynamics software to show the characteristics of the flow over a reverse delta wing.

The increase of the worldwide greenhouse gas emissions is a growing danger. The global carbon emission from fossil fuels has raised through the 10,000 million metric tonnes by 2014 [36]. The aviation sector has its part in this global greenhouse effect, and several measures were taken to keep under control such as Kyoto Protocol. An approach to reduce this effect is to enhance the aerodynamic characteristics of the aircrafts. Hence, to improve the fuel efficiency of the aerial vehicles. Improving the aerodynamic performance of the manned or unmanned aerial vehicles via active and passive flow control technologies is a key point for this issue. Leading-edge undulations and plasma vortex generators were started to be established in the 1990s, and they continue to be developed. Leading-edge undulations have a sufficient agreement in the improving the aerodynamic efficiency of the post-stall regime. However, those passive flow control devices do not have a significant effect in the pre-stall regime, especially in the cruise conditions. On the other hand, plasma vortex generators are known for their beneficial effects in the drag reduction and the mixing enhancement in the boundary layer for the lift improvement. In this project, the plasma vortex generators will be used as leading-edge vortex generator devices to delay stall, to improve post-stall aerodynamic efficiency and to maintain the pre-stall lift-drag ratio. After the conducted experiments, it was shown that the plasma vortex generators provided more promising results than the undulations in low Reynolds numbers, and capable of being used in higher Reynolds numbers.

An investigation of the flow field around a double-delta wing was carried out at the Aerodynamics Analysis and Design
Laboratory’s (AADL) water tunnel facility. The model has a sharp leading edge and 76=40
configuration, which is representative of current
fighter aircraft. The experiments were conducted at Reynolds numbers of 10,000 and 15,000. The dye injection technique was utilized to analyze
the flow.Acomparison of the vortical flow structure and the bursting point locations was done for a Reynolds number of 15,000 and found to be
in reasonable agreement. The major objective of this experiment was to study the effect of the angle of attack (AOA) and Reynolds number on
the vortical flow structure. In this flow visualization result, the dominant feature of the AOA and Reynolds number effect on the vortical flow
structure over the double-delta wing has been observed. With increasing Reynolds number, the distance between the vortex trajectory and the
model surface becomes smaller. As the AOA increases, the intertwining or coiling up features cannot be seen and the vortex bursting point
locations move upstream. The distance between the vortical streakline and the model surface becomes larger with increasing AOA. The
crossover point of the main and strake wing vortices also depends on the AOA.

The unsteady flow field around oscillating OA-209
airfoil at a Reynolds number of 3.5×105 were investigated. Three
different reduced frequencies were tested in order to see how it
affects the hysteresis loop of an airfoil. At a reduced frequency of
0.05 the deep dynamic stall phenomenon was observed. Lift
overshooting was observed as a result of dynamic stall vortex (DSV)
shedding. Further investigation was carried out to find out the cause
of DSV formation and shedding over airfoil. Particle image
velocimetry (PIV) and CFD tools were used and it was found out that
dynamic stall separation (DSS), which is separated from leading edge
separation (LES) and trailing edge separation (TES), triggered the
dynamic stall vortex (DSV).

This paper describes the experimental and computational work carried at the university facility. The experimental winglet model made of Medium Density Fiber (MDF) board while computational model was created using CATIA V6 software adopting NACA 0012 airfoil. The experimental work was performed in open loop Low Speed Subsonic Wind Tunnel of test cross section of 0.6 m x 0.6 m x 2 m at different Reynold Numbers suited for MAV and at different angles of attack. The CFD simulation was performed using 3-D structured hexahedral mesh generated in ICEM CFD software. In viscid Euler and laminar RANS k-ω turbulent model with finite volume method were used. The various aerodynamic characteristics were measured and analyzed. The results show that the corrugation winglet produces better lift and better drag reduction compared to other two configurations tested and hence the significant performance (L/D Ratio) improvement nearly 2.5 to 3 times of plain wing is obtained.

A six component force balance was used to obtain the C L for a reverse delta wing and a delta wing for comparison purposes. A simulation of the streamlines, velocity vectors and surface pressure contours was carried out using CFD software to show the characteristics of the flow over a reverse delta wing.

This paper describes the results of an experimental investigation of the wake region of a model wind turbine rotor with tip injection. The experiments are conducted by placing a specially designed three-bladed horizontal axis wind turbine rotor at the exit of an open-jet wind tunnel facility. The rotor blades are non-linearly twisted and tapered with NREL S826 airfoil profile all along the span. The nacelle, hub and the blades are specifically designed to allow pressurized air to pass through and get injected from the blade tips while the rotor is rotating. The setup is instrumented with a torque meter and a load cell to allow measurements of torque, thrust, side loads and the rotational speed of the rotor and it also allows for Particle Image Velocimetry (PIV) measurements as well as Constant Temperature Anemometry (CTA) measurements downstream of the turbine rotor. This paper will present the results of CTA measurements both for baseline (no-injection) as well as for tip injection cases focusing on the effects of tip injection on the wake flow field. Results show that, when there is injection, obtained characteristics have remarkable differences compared to the baseline case both for the load data showing an increase in power and thrust coefficients for TSR values starting near maximum CP condition up to higher TSR levels as well as for the wake characteristics showing a tip flow region that is radially pushed outwards and with increased levels of turbulence occupying wider areas compared to the baseline case. Within the wake zone covering a region up to one diameter downstream, it’s observed that the boundary between the wake and the freestream gets wider, more diffused and turbulence levels increase due to tip injection. Final paper will present more CTA data within the wake up to two diameters downstream as well as some 2D PIV data on streamwise planes within the wake focusing on the effects of tip injection on the tip vortex trajectory and turbulence levels.

Effects of the sideslip angle on the wing-kinematics and the aerodynamic forces and moments generated by the locusts were investigated in a low-speed wind tunnel. Three locusts (Schistocerca americana) were tested at tunnel speeds of 0,
2, and 4 m/s at three body angles of attack (0, 3 and 7 degrees) and three sideslip angles (0, -10 and -20 degrees) at each angle of attack. A sensitive custom-built microbalance was used to measure the forces and moments. Simultaneous high -
speed videos were taken to record the wing kinematics. Results in the form of normalized stroke-averaged forces and moments are presented and a correlation with the wing kinematics has been investigated.

A systematic investigation into the design and simulation of flow parameters in a closed-loop wind tunnel was carried out using Computational Fluid Dynamics (CFD). The analytical model for estimating pressure losses were directed as input boundary conditions. Full-scale model of the entire wind tunnel was considered instead of the conventional approach, in which only test section flow is simulated. This allowed for optimisation of flow quality not only in the test section but also the flow in the entire circuit. Analysis of the guide vane configurations showed that test section flow quality was more affected by flow conditions in upstream than downstream sections. Hence, special attention must be given while designing the vanes at upstream turns particularly corners in line with the test section. Validation of the test section with block model showed that CFD was able to replicate wind tunnel measurements of velocity, turbulence intensity and pressure coefficient with error below 10%.

Bluff bodies may assume arbitrary attitudes in a flow, causing aerodynamic loads that are sensitive to attitude. The Continuous Rotation technique obtains 6-component loads on bluff bodies with 1-degree azimuth resolution about selected axes at a rate of 1 revolution per minute. The load coefficient variation is Fourier transformed and the resulting complex series is truncated in order to obtain rapidly computable, analytical formulae. The method is applied to bluff body shapes including cylinders, a cuboid, a flat plate and a porous box. A cylinder whose length is equal to its diameter, is used to show that rate effects, hysteresis, vortex shedding and other unsteady aerodynamic phenomena are negligible below 10 revolutions per minute. Approaches to generalize the aerodynamic loads on yawed finite cylinders of various aspect ratio are studied. The reasons for differences in aerodynamic load behavior between 2 cylinder models, are analyzed. To complement experiments, the ROTCFD unsteady Navier Stokes code is used to perform diagnostic computations. Methods to generalize the predictions are explored. Maps of the leading coefficients of the Fourier series of each load component over the aspect ratios space, are interpolated. The interpolation varies sharply for aspect ratios between 0.5 to 1. The variation is more gradual beyond aspect ratio 2. By aspect ratio 4, a 'high aspect ratio' limit appears to be reached.

Particle Image Velocimetry was used in a low speed wind tunnel to investigate the vortex structures of a slender reverse delta wing at various angles of attack and roll. This work investigates the characteristics of the vortices generated downstream in a plane perpendicular to the free stream direction and their dependence on angles of attack and roll angles at chord based Reynolds number.

In this paper, variation of static torque and drag coefficient of an S-Shaped Savonius rotor have been studied for different number of rotor blades. This has been done by measuring the pressure distribution on the blade surfaces for different rotor angles. All the experiments have been carried out at Reynolds number 1.07x105 in a uniform flow jet produced by an open circuit wind tunnel. Experimental results indicate that both of the drag force and hence the torque change frequently with rotor angle. It is also found that as the increase of the number of rotor blade increases the starting torque of rotor increases but there is no significant increase in net torque output. It is observed that, torque distribution becomes more uniform as the number of blade of rotor increases.

Flight pressure and heat flux data have been compared to angle-of-attack-and yaw-dependent computational-fluid-dynamics results for pressure distribution as well as laminar and turbulent heat-transfer results at three time points in the ascent trajectory. Computed pressures, normalized by freestream pressure, were interpolated to the flight Mach numbers at each time point throughout the ascent and descent trajectories, and angle of attack and yaw were estimated from measured pressure by determining the combination minimizing the difference between the measured and computed pressures. The resulting vehicle attitude was compared to the vehicle attitude derived from inertial measurement unit results from the flight. The two methods showed excellent agreement for the entirety of the ascent and reentry portions of the trajectory. A similar normalization of the laminar and turbulent computational-fluid-dynamics heat transfer results into Stanton number distributions was compared to flight heat transfer measurements, and transition times at a given location were inferred. Computational heat conduction analysis verified assumptions in the calculation of heat flux from temperature. Nomenclature c p = specific heat capacity, constant pressure, J∕…kg ⋅ k† M = Mach number p = pressure, Pa _ q = heat flux, W∕m 2 Re = Reynolds number St = Stanton number T = temperature, K t = time, s u = velocity, m∕s x = streamwise distance, m α = angle of attack, deg β = yaw, deg ρ = density, kg∕m 3 ϕ = angular location on vehicle surface, deg Subscripts CFD = derived from computational solution CL = centerline e = boundary-layer edge conditions F = derived from flight measurement FFT = fast Fourier transform IMU = recorded by the inertial measurement unit LE = leading edge RMS = derived from root-mean-squared error minimization between computational fluid dynamics and flight tr = at transition onset w = wall conditions 0 = stagnation conditions ∞ = freestream conditions

Aerodynamic and inertial forces and corresponding kinematics of flapping wings of locusts,
Schistocerca americana, were investigated in a low-speed wind tunnel. The experimental setup
included live locusts mounted on microbalance synchronized with a high-speed video system.
Simultaneous measurements of wing kinematics and forces were carried out on three locusts at 7°
angle of attack and velocities of 0ms−1 and 4ms−1. Time variations of flapping and pitching angles
exhibit similar patterns in fore- and hindwings and among the animals. Significant tip to root
variations in pitching angle are found in both wings. The locusts have much larger flapping and
pitching amplitudes in still air causing larger oscillations in inertial forces. Inertial forces are added to
the lift and thrust on one part of the stroke, resulting in higher reaction forces and subtracted on the
other part. Plots of the lift demonstrate similar trends with and without the wind. The global maxima
and peak-to-peak amplitudes in lift are about the same in both tests. However, local minima are
significantly lower in still air, resulting in much smaller stroke-averaged lift. Amplitudes of thrust
force oscillations are much higher in still air; consequently, the stroke-averaged thrust is higher
compared to the non-zero freestream velocity case.

Aerodynamic and inertial forces and corresponding kinematics of flapping wings of locusts, Schistocerca Americana, were investigated in a low-speed wind tunnel. The experimental setup included live locusts mounted on microbalance synchronized with a high-speed video system. Simultaneous measurements of
wing kinematics and forces were carried out on the three locusts at zero velocity and at 4 m/s and 7° angle of angle of attack. Time variations of flapping and pitching angles exhibit similar patterns in fore- and hindwings and among animals. Significant tip to root variations in pitching angle are found in both wings. The locusts have much larger flapping and pitching amplitudes in still air causing larger oscillations in inertial forces. Inertial forces are added to the lift and thrust on one part of the stroke, resulting in higher reaction forces and subtracted on the other part. Plots of the lift demonstrate similar trends with and without the wind. The global maxima and peak-to-peak amplitudes in lift
are about the same in both tests. However, local minima are significantly lower at zero freestream, resulting in much smaller stroke-averaged lift. Amplitudes of thrust force oscillations are much higher at zero freestream, consequently, the stroke-averaged thrust is much higher compared to the non-zero freestream velocity case.

The calibration tests based on the ISSI BF 400 paint have been shown and the obstacles were addressed making these successful measurements. There are slight changes in the calibration curves according to different reference temperature and pressure values. The effects of these changes on the interpretation of delta wing flow are well discussed in this paper.

Maximum drag achievement is the main purpose in parachute design. Common assumptions in parachutes simulation are simplified inflated canopy geometry and impermeability of parachute canopy. In this research, these assumptions result in evaluation reduction of numerical simulation. By considering the permeability and its effects on flow domain, accuracy of numerical simulation increases. Coefficients of Darcy's modified equation were resulted from experiments applied for numerical simulation of some simplified canopies. Assuming permeability for the canopy, the drag coefficient showed a 24 percent decrease compared to the impermeable one. By experimental tests, drag of the permeable and impermeable canopy was measured in low speed wind tunnel test. Geometries of inflated canopies in wind tunnel were used in numerical simulation and then the numerical results were evaluated by the experimental ones. Comparison between solid canopy and permeable one, showed significant differences in the results, especially in streamlines, pressure distribution and drag coefficient. These differences highlight considering permeability assumption in numerical simulation of parachute canopies.

Unilamellar liposomes with incorporated hapten–phospholipid conjugates were proposed as models of polyvalent antigens with migrating determinants for quantitative analysis of their interaction with antibodies. The monovalent pesticide atrazine was used as a model antigen. For its incorporation into the lipid bilayer, the atrazine carboxylated derivative was conjugated with dimyristoylphosphatidylethanolamine (DMPE). Unilamellar liposomes were prepared with dimyristoylphosphatidylcholine/atrazine–DMPE at molar ratios of

ABSTRACT A transonic natural-laminar-flow wing design procedure has been set up, integrating a parametric geometry model with several analysis tools. A direct design strategy has been applied and three levels of aerodynamic analysis have been used: a full-potential method (with which to rapidly iterate to obtain the target pressure distribution), an Euler solution coupled with a boundary-layer solver and a semi-empirical stability analysis method (for an intermediate evaluation that is able to include the laminar flow extension), and a full Navier-Stokes analysis with fixed transition (for a final verification of the design quality). A transonic natural-laminar-flow wing suitable for business aviation able to perform at least 40% of laminar flow in cruise with acceptable wave drag at Mach 0.78 has been designed and tested at flight Reynolds numbers. The laminar flow has been verified by infrared cameras.

In general, wind turbine blades operate in subsonic regions along with turbulent flow conditions. This paper deals with noise due to turbulent inflow around the blades based on Amiet's and Lowson's model formulation .This mechanism is predicted using numerical simulation for different observer azimuth positions and various rotor speeds of turbine. The influence of turbulence intensity, length scales on the sound pressure levels is illustrated. The effect of blade geometry characteristics on sound pressure levels is implemented. The influence on power production from wind turbine due to the reduction in sound power levels are correlated with the modified blade geometry. The total noise from turbulent boundary layer trailing edge and turbulent inflow mechanisms are compared for three turbines of sizes, 350kW, 2MW & 3MW. Overall sound pressure levels are dominated by the inflow turbulence mechanism. The amplitude modulation is calculated using the sound pressure levels obtained and illustrated for predominantly low and medium range frequencies in octave spectra.

Experimental testes were carried out to investigate the flow near a flap side-edge with attached porous plate as a potential noise reduction device. The experiments were performed over a wing to simplify the installation. The objective was to perform a first approximation to the flow field produced by the addition of porosity of different sizes. Specifically, the changes in tip vortex were studied. Baseline case and plates of 10mm, 20mm, 30mm, 50mm and 70mm placed in the bottom side of the wing tip were tested. Also another configuration was tested, with a plate of 25mm roughly placed in the attachment line of the vortex. Aerodynamic experiments were performed to evaluate the changes in the lift, drag and pitching moment coefficients. Hot wire anemometry measurements were performed in perpendicular planes to the direction of the undisturbed flow. Problems in the calibration were detected, that corrupts all the values. Nevertheless, first results about the behavior of the flow field were achieved. The addition of the porous plate transforms the structure of the vortex by dramatically decreasing the axial velocity in the core center transforming the tip vortex from a jet vortex to a wake vortex. Vorticity was also changed with a noticeable decrease. From 30mm porous strip, this decrease becomes asymptotic. There are some structures that are not completely explained. The attachment line plate and the 10mm plate cases present a complex structure, which seems due to various vortices. The 50mm and 70mm plates show a merging between the wake and the vortex. The addition of the porous plate weakens the mechanisms of noise production. It is judged that this will lead to a noise reduction.

An extensive experimental investigation is conducted to study the effect of canard position UHODWLYHHWRRWKHHIXVHODJHHUHIHUHQFHHOLQHHRQQWKHHDHURG\QDPLFFIRUFHVVRIIDD¿JKWHUUW\SHHFRQ¿JX-UDWLRQQPRGHO$HURG\QDPLFFIRUFHVVDWWGLIIHUHQWWÀLJKWWFRQGLWLRQVVDUHHPHDVXUHGGLQQDDVXEVRQLFF wind tunnel. The wing and the canard have triquetrous shapes. Experiments are conducted DWW5H\QROGVVQXPEHUURIIDQGGDWWWRRGHJUHHHDQJOHVVRIIDWWDFN7KHHUHVXOWVVVKRZZWKDWW FDQDUGGLQFUHDVHVVWKHHOLIWWDQGGGUDJJIRUFHVVZKLOHHLWWGHFUHDVHVVWKHHVWDWLFFVWDELOLW\\RIIWKHHPRGHO The canard at its up position

The HIFiRE-5 test article was an elliptic cone with a 2.5-mm nose radius and 2:1 aspect ratio and a 7-degree minor-axis half-angle. The vehicle was flown in April 2012. The upper stage of the sounding rocket failed to ignite, resulting in a peak Mach number of about 3 instead of the target of 7. Flight heat flux and pressure data (reduced from almost 300 thermocouples and 50 pressure transducers) have been compared to α-and β-dependent CFD results for pressure distribution, as well as laminar and turbulent heat-transfer results. Computations were performed at three time points in the ascent trajectory. At each time point, five values each of angle of attack and yaw, ranging from -5.0° to 5.0°, were computed. CFD pressures, normalized with p ∞ , were interpolated to the flight Mach numbers at specified times throughout the ascent and descent trajectories. At each flight time, α and β were estimated from measured pressure by determining the α-β combination that minimized the RMS difference between the measured and computed pressures. The vehicle attitude, as determined from measured pressure, was compared to the vehicle attitude derived from Inertial Measurement Unit (IMU) results for α and β from the flight. The two methods showed excellent agreement for the entirety of the ascent and reentry portions of the trajectory. A similar normalization of the laminar and turbulent heat transfer CFD results with St was compared to flight heat transfer measurements, and transition locations were inferred. Finally, a computational heat conduction analysis was made to verify assumptions inherent in the calculation of heat flux from temperature.

The HIFiRE-5 test article was an elliptic cone with a 2.5-mm nose radius and 2:1 aspect ratio and a 7-degree minor-axis half-angle. The vehicle was flown in April 2012. The upper stage of the sounding rocket failed to ignite, resulting in a peak Mach number of about 3 instead of the target of 7. Flight heat flux and pressure data (reduced from almost 300 thermocouples and 50 pressure transducers) have been compared to α-and β-dependent CFD results for pressure distribution, as well as laminar and turbulent heat-transfer results. Computations were performed at three time points in the ascent trajectory. At each time point, five values each of angle of attack and yaw, ranging from-5.0° to 5.0°, were computed. CFD pressures, normalized with p ∞ , were interpolated to the flight Mach numbers at specified times throughout the ascent and descent trajectories. At each flight time, α and β were estimated from measured pressure by determining the α-β combination that minimized the RMS difference between the measured and computed pressures. The vehicle attitude, as determined from measured pressure, was compared to the vehicle attitude derived from Inertial Measurement Unit (IMU) results for α and β from the flight. The two methods showed excellent agreement for the entirety of the ascent and reentry portions of the trajectory. A similar normalization of the laminar and turbulent heat transfer CFD results with St was compared to flight heat transfer measurements, and transition locations were inferred. Finally, a computational heat conduction analysis was made to verify assumptions inherent in the calculation of heat flux from temperature.

Experimental tests were performed to investigate the flow near a flap side-edge with attached porous plate as a potential noise reduction device. The experiments were performed using a single wing to simplify the installation. The objective was to perform a first approximation to the flow field produced by the addition of porosity of different sizes. Specifically, the changes in tip vortex were studied. Baseline case and plates of 10mm, 20mm, 30mm, 50mm and 70mm placed in the bottom side of the wing tip were tested. Also, another configuration was tested, with a plate of 25mm roughly placed in the attachment line of the vortex. Aerodynamic experiments were performed to evaluate the changes in the lift, drag and pitching moment coefficients. Hot wire anemometry measurements were performed in perpendicular planes to the direction of the undisturbed flow. Problems in the calibration were detected that corrupts all the values. Nevertheless, first results about the behavior of the flow field were achieved. The addition of the porous plate transforms the structure of the vortex by dramatically decreasing the axial velocity in the core center transforming the tip vortex from a jet vortex to a wake vortex. Vorticity was also changed with a noticeable decrease. From 30mm porous strip, this decrease becomes asymptotic. There are some structures that are not fully explained. The attachment line plate and the 10mm plate cases present a complex structure, which seems due to various vortices. The 50mm and 70mm plates show a merging between the wake and the vortex. The addition of the porous plate weakens the mechanisms of noise production as it reduces vorticity, core flow velocity and displace it away to the wing tip.

This paper describes the experimental and computational work carried at the university facility. The experimental winglet model made of Medium Density Fiber (MDF) board while computational model was created using CATIA V6 software adopting NACA 0012 airfoil. The experimental work was performed in open loop Low Speed Subsonic Wind Tunnel of test cross section of 0.6 m x 0.6 m x 2 m at different Reynold Numbers suited for MAV and at different angles of attack. The CFD simulation was performed using 3-D structured hexahedral mesh generated in ICEM CFD software. Inviscid Euler and laminar RANS k-ω turbulent model with finite volume method were used. The various aerodynamic characteristics were measured and analyzed. The result shows that the corrugation winglet produces better lift and better drag reduction compared to other two configurations tested and hence the significant performance improvement is obtained. 1.0 Introduction Aerodynamic performance of Micro Arial Vehicles (MAV) is highly depends on the induced drag caused by the wingtip vortices. Winglets are used to reduce the induced drag there by enhancing efficiency, range and endurance of the MAV. Most of the natural flyers (like birds and insects) wings are not profiled airfoil but having well-structured corrugation/zigzag configuration. The aircraft manufacturers are under immense pressure to improve efficiency due to rising operating cost and environment issues and have initiated to develop innovative induced drag reduction technology. One of the methods to reduce this drag is to use wingtip devices called winglets. The first concept of winglet known as end plat was developed and patented by F.W.Lancaster in 1897. A winglet using advanced airfoil concept to reduce induced drag was designed and tested by R.Whitecomb between

Particle Image Velocimetry was used in a low speed wind tunnel to investigate the vortex structures of a slender reverse delta wing at various angles of attack and roll. This work investigates the characteristics of the vortices generated downstream in a plane perpendicular to the free stream direction and their dependence on angles of attack and roll angles at chord based Reynolds number.

In nature, many fish aggregate into large groups or schools for protection against predators, for social interactions and to save energy during migrations. Regardless of their prime motivation, fish experience three-dimensional flow interactions amongst themselves that can improve or hamper swimming performance and give rise to fluid-mediated forces between individuals. To date, the unsteady, three-dimensional flow interactions among schooling fish remains relatively unexplored. In order to study these interactions, the caudal fins of two interacting fish are idealized as two finite span pitching wings arranged in mixtures of canonical in-line and side-by-side arrangements. The forces and moments acting on the wings in the streamwise and cross-stream directions are quantified as the arrangement and the phase delay between the wings is altered. Particle image velocimetry is employed to characterize the flow physics during high efficiency locomotion. Finally, the forces and flowfields of two-dimensional pitching wings are compared with three-dimensional wings to distinguish how three-dimensionality alters the flow interactions in schools of fish.

Experimental tests were performed to investigate the flow near a flap side-edge with attached porous plate as a potential noise reduction device. The experiments were performed using a single wing to simplify the installation. The objective was to perform a first approximation to the flow field produced by the addition of porosity of different sizes. Specifically, the changes in tip vortex were studied. Baseline case and plates of 10mm, 20mm, 30mm, 50mm and 70mm placed in the bottom side of the wing tip were tested. Also, another configuration was tested, with a plate of 25mm roughly placed in the attachment line of the vortex. Aerodynamic experiments were performed to evaluate the changes in the lift, drag and pitching moment coefficients. Hot wire anemometry measurements were performed in perpendicular planes to the direction of the undisturbed flow. Problems in the calibration were detected that corrupts all the values. Nevertheless, first results about the behavior of the flow f...