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In this project work, experiments are carried out in low-speed subsonic wind tunnel and 30-port multitube manometer used to study the aerodynamics characteristic of cambered airfoil model NACA5520 with different velocity range and various angle of attack. Experiments were made to investigate the pressure distribution over the airfoil under velocity range from 5m/s, 10m/s, 15 m/s, and 20 m/s and various x/s locations as 0.25, 0.5, and 0.75. The pressure measurements were taken using a multi-tube manometer and pitot-static tube over the cambered airfoil at different angles of attack. Hence, the distribution of the pressure coefficient (Cp) is calculated for the different operating conditions on the tapered blade. Also, the effect of pressure distribution over the tapered blade at different velocities and AOA is studied in spanwise location of the blade. [ABSTRACT FROM AUTHOR]

Copyright of Journal of Polytechnic is the property of Journal of Polytechnic and its content may not be copied or emailed to multiple sites or posted to a listserv without the copyright holder's express written permission. However, users may print, download, or email articles for individual use. This abstract may be abridged. No warranty is given about the accuracy of the copy. Users should refer to the original published version of the material for the full abstract. (Copyright applies to all Abstracts.)

Featured Application: This study is meaningful for the design and safe flight of the wing-in-ground craft. Wing-in-ground craft often encounter waves when flying over the sea surface, and the ground effect is more complicated than that of flat ground. Therefore, the aerodynamic characteristics of the NACA 4412 airfoil in proximity to wavy ground for a wide range of angles of attack is studied by solving the Reynolds Averaged Navier–Stokes equations. The validation of the numerical method is carried out by comparing it with the experimental data. The results show that the aerodynamic coefficients will fluctuate periodically when the airfoil moves over wavy ground at a small ride height. Except for the angle of attack of 0°, the fluctuation trend of aerodynamic coefficients at other angles of attack is the same. The analysis of aerodynamic fluctuation amplitude found that the medium angle of attack should be selected as the design cruise angle of attack for wing-in-ground craft. The time-averaged aerodynamic coefficients in the case of wavy ground are almost the same as those of flat ground. Hence, wavy ground mainly causes a fluctuation in aerodynamic coefficients. Considering the difference between aerodynamic coefficients at the angle of attack of 0° and at other angles of attack, the flow field structure at an angle of attack of 0° and 4° is analyzed. The results reveal the aerodynamic characteristics of the airfoil moving over wavy ground, which gives a deeper understanding of the ground effect in the conditions of wavy surface/ground. This has a certain guiding significance for the design of wing-in-ground craft. [ABSTRACT FROM AUTHOR]

The research entails a finite-element design of a 3d-autopilot missile synthesis of multiple objective controls by solving the inequalities encountered while using a linear matrix. The autopilot design analyses the performance methodologies for a group of finite models that are further linearised at different aerodynamic roll angles within the flight coverage for harnessing the uncertainties that are encountered confined within the high angle-of-attack region. Multiple simulation results are devised at various roll angle conditions to satisfy the performance of the controls. [ABSTRACT FROM AUTHOR]

Spectral analysis was performed on the time series data computed from pressure measurements on the New MEXICO (Model Rotor Experiments under Controlled Conditions) rotor in standstill conditions. As a priori, 3D airfoil polars were recreated from standstill measurements and compared against 2D airfoil polars and flat plate theory results to verify the measurements. The spectral analysis revealed the presence of dominant shedding frequencies for certain ranges of the geometric angle of attack. Two dominant shedding modes were identified: One was associated with bluff body vortex shedding, and the other was associated with low Strouhal number shedding. No dominant shedding frequencies were observed for angles of attack beyond 50°. The research improves on our current understanding of the unsteady nature of the stall regime, along with providing insight into the existence of vortex‐induced vibrations on a wind turbine in standstill conditions. [ABSTRACT FROM AUTHOR]

With distributed propulsion and electric vertical take-off and landing aircraft on the rise, fast and accurate methods to simulate propeller slipstreams and their interaction with aircraft components are needed. In this work, we compare results obtained with a filament-based free wake panel method to experimental and previously validated numerical data. In particular, we study a propeller-wing configuration at zero angle of attack and the aerodynamics of the blade-resolved slipstream interaction with the wing. We use a prescribed wake on the wing and a free wake on the propeller, which greatly accelerate the computations. Results indicate that, while forces are overpredicted due to the inviscid nature of the panel method, the free wake is able to capture the slipstream deformation and shearing with remarkable success. We find that a filament-based free wake panel method can be a useful tool for propeller-wing interaction in preliminary aircraft design. [ABSTRACT FROM AUTHOR]

Purpose The purpose of this paper is to propose a high angle of attack short landing model for switched polytopic systems as well as to derive an equation for fluidic thrust vector deflection angle based on pressure to reduce the velocity during the landing phase of flight.Design/methodology/approach In this paper, robust control algorithm is proposed for a non-linear high angle of attack aircraft under the effects of non-linearities, tottering hysteresis, irregular and wing rock atmosphere. High angle of attack short landing flight under asynchronous switching is attained by using the robust controller method. Lyapunov function and the average dwell time scheme is used for obtaining the switched polytopic scheme. The asynchronous switching and loss of data are controlled asymptotically. The velocity of aircraft has been lucratively reduced during the landing phase of flight by using the robust controller technique.Findings The proposed algorithm based on robust controller including the effects of non-linearities guarantee the successful reduction of velocity for high angle of attack switched polytopic systems.Practical implications As the landing phase of an aircraft is one of the complicated stage, this algorithm plays a vital role in stable and short landing under the condition of high angle of attack (AOA).Originality/value In this paper, not only the velocity of flight has been reduced, but also the high angle of attack has been attained during the landing phase, because of which the duration of landing has been reduced as well, while in most of the previous research, it is based on low angle of attack and long landing duration. [ABSTRACT FROM AUTHOR]

A novel synthetic air data estimation method without using air data sensors is presented, and the method only relies on the information from the Navigation System (NS) and Flight Control System (FCS). The aircraft's aerodynamic model is also required to make a connection between the FCS control parameters and the NS measurements. The airspeed, angle of attack and sideslip, angular velocity and wind speed are defined as state vectors, and state equations are established through the aircraft's aerodynamic model and dynamics. Linear velocity and angular velocity provided by the navigation system are considered as the measurement vector. To deal with variable wind fields, a novel Initialised Three-step Extended Kalman Filter (ITEKF), which considers the wind speed as an unknown input, is developed to track the variation of wind speed. Simulation results based on a Generic Hypersonic Vehicle (GHV) model are presented and compared with an existing method. Factors affecting the method's accuracy include the navigation system accuracy and the aerodynamic model error, are also discussed. [ABSTRACT FROM AUTHOR]

Abstract This research presents a nonlinear gain-scheduled flight controller design method via a stable manifold theory in order to handle the nonlinearities of the F-18 High Alpha Research Vehicle due to the change in the aerodynamic characteristics at different angles of attack and the airspeed variation. The designed longitudinal flight control system consists of a nonlinear gain-scheduled stabilization augmentation system which is designed using the stable manifold method, and a linear gain-scheduled control augmentation system which consists of proportional and integral gains. The nonlinear longitudinal flight controller is verified in a 6 degree-of-freedom simulator. [ABSTRACT FROM AUTHOR]

In this paper, based on the thermo dynamic design of the supercritical carbon dioxide (sCO2) centrifugal compressor, the design idea of the flow passage geometries and the method to improve the performance of the sCO2 centrifugal compressor are discussed. With the help ofcommercial software ANSYS CFX, the influence of the shape of the leading edge and trailing edge is studied, and the elliptical leading edge makes the pressure ratio 10.30% higher and the efficiency 3.95% higher than the square leading edge. By changing the forward-swept angle and backward-swept angle of the leading edge, the effects of aerodynamic swept shape in sCO2 centrifugal compressor are discussed. The effect of the gap between the impeller blade and diffuser blade is discussed, and the 10 mm gap makes the performance best. The pressure ratio is increased by 2.5% compared with the original design, while at the same time the efficiency is slightly improved. In summary, based on thermal design of the sCO2 centrifugal compressor, the effects of different flow geometries are analyzed in detail. [ABSTRACT FROM AUTHOR]

Over the last decades, there has been great interest in understanding the aerodynamics of flapping flight and development of flapping wing Micro Air Vehicles (FWMAVs). The camber deformation and twisting has been demonstrated quantitatively in a number of insects, but making artificial wings that mimic those features is a challenge. This paper reports the development and characterization of artificial wings that can reproduce camber and twisting deformations. By replacing the elastic material at the wing root vein, the root vein would bend upward and inward generating an angle of attack, camber, and twisting deformations while the wing was flapping due to the aerodynamic forces acting on the wing. The flapping wing apparatus was employed to study the flexible wing kinematics and aerodynamics of real scale insect wings. Multidisciplinary experiments were conducted to provide the natural frequency, the force production, three-dimensional wing kinematics, and the effects of wing flexibility experienced by the flexible wings. The results have shown that the present artificial wing was able to mimic the two important features of insect wings: twisting and camber generation. From the force measurement, it is found that the wing with the uniform deformation showed the higher lift/power generation in the flapping wing system. The present developed artificial wing suggests a new guideline for the bio-inspired wing of the FWMAV. [ABSTRACT FROM AUTHOR]

This paper presents an experimental investigation on using FOLDing wingtips sERving as cONtrol effectorS (FOLDERONS) for a mini Unmanned Aerial Vehicle (UAV). A representative off-the-shelf mini-UAV with a conventional configuration was selected. The main theme of this paper is to utilise FOLDERONS as a control effector (mainly in roll) to augment the control authority of conventional control surfaces. Furthermore, the impact of actuation rate on the effectiveness of FOLDERONS is assessed. The paper describes the preliminary and detailed design and sizing of the morphing wing. In addition, the manufacturing of the wing system and its integration with the UAV are addressed. Wind-tunnel testing in the RJ Mitchell wind-tunnel at the University of Southampton was performed. Both static (straight and sideslip) and dynamic (straight flight) tests are conducted at a range of airspeeds and Angles Of Attack (AOAs). The impact of folding wingtips on the lateral and directional stability is analysed. The main finding of this paper is that FOLDERONS are effective (especially at large dynamic pressure and AOAs) in controlling the lateral and directional stability. Finally, this study shows that FOLDERONS cannot fully replace conventional ailerons especially at low dynamic pressures, and their strong dependence on the AOA makes them prone to a roll reversal phenomena when the wing (and FOLDERONS) is operating at negative AOAs. [ABSTRACT FROM AUTHOR]

Purpose This paper aims to investigate the aerodynamic characteristics as well as static stability of wing-in-ground effect aircraft. The effect of geometrical characteristics, namely, twist angle, dihedral angle, sweep angle and taper ratio are examined.Design/methodology/approach A three-dimensional computational fluid dynamic code is developed to investigate the aerodynamic characteristics of the effect. The turbulent model is utilized for characterization of flow over wing surface.Findings The numerical results show that the maximum change of the drag coefficient depends on the angle of attack, twist angle and ground clearance, in a decreasing order. Also, it is found that the lift coefficient increases as the ground clearance, twist angle and dihedral angle decrease. On the other hand, the sweep angle does not have a significant effect on the lift coefficient for the considered wing section and Reynolds number. Also, as the aerodynamic characteristics increase, the taper ratio befits in trailing state.Practical implications To design an aircraft, the effect of each design parameter needs to be estimated. For this purpose, the sensitivity analysis is used. In this paper, the influence of all parameter against each other including ground clearance, angle of attack, twist angle, dihedral angle and sweep angle for the NACA 6409 are investigated.Originality/value As a summary, the contribution of this paper is to predict the aerodynamic performance for the cruise condition. In this study, the sensitivity of the design parameter on aerodynamic performance can be estimated and the effect of geometrical characteristics has been investigated in detail. Also, the best lift to drag coefficient for the NACA 6409 wing section specifies and two types of taper ratios in ground effect are compared. [ABSTRACT FROM AUTHOR]

This paper investigates optimal flight of aerodynamically controlled missiles in the terminal phase with both angle of attack and bank angle as control inputs. Practical constraints on impact angle and dynamic pressure are all considered. To efficiently and reliably solve such a highly nonlinear (nonconvex) problem, we first present how to obtain its convex relaxation and then propose a regularization technique which is critical to guarantee the exactness of this relaxation. Numerical examples are given to illustrate the effectiveness of the proposed method. [ABSTRACT FROM PUBLISHER]

Wing stiffness is very crucial in augmenting aerodynamic forces in flapping wing flyers. In this work, the effect of wing deformation was studied using three-dimensional numerical analysis (two-way fluid structure interaction), coupling the flow solver (FLUENT) and the structural (ABAQUS) solver via the MpCCI platform. Three different degrees of bending stiffness corresponding to rigid, flexible, and highly flexible case wings were investigated. Moreover, the wings were tested for both low Reynolds number (R ¼ 9,000) and high Reynolds number (R ¼ 40,000), at a flapping frequency of 9 Hz corresponding to an angle of attack (AoA) ranging from α ¼ 0 to 50°. The results of mean aerodynamic lift and drag coefficients showed good agreement between numerical and experimental findings. Also, the time-averaged lift-to-drag ratio reveals that the highly flexible wing exhibited the best overall aerodynamic performance when compared to the rigid and flexible wing. [ABSTRACT FROM AUTHOR]

In this paper, the flow around a rectangular cylinder with aspect ratio 5:1 is studied at three attack angles by means of both LES and URANS simulations. Despite its geometric simplicity, such a rectangular shape is characterized by many aerodynamic phenomena typical of reattached flows around bluff bodies and has been recently widely studied as a prototype of flows encountered around shapes of technical interest like, for example, bridge deck sections. The present study is aimed at providing insight in the modifications occurring in the flow at small incidence angles, by comparing the accuracy of the two aforementioned simulation strategies in reproducing them. Some recurring biases observed in URANS simulations are illustrated and their origin discussed. Results are presented in terms of flow bulk parameters and pressure distributions and systematic comparison with available experimental data is provided. Finally, the effects of small incidence angles on along span flow correlations are investigated and Covariance Proper Transformation is used in order to further characterize the flow dynamic behavior. [ABSTRACT FROM AUTHOR]

An experimental double multi-modal flutter instability was recorded during wind tunnel tests on a full bridge aero elastic model. The analysis of this instability is presented in this paper, and it is studied with a multi-modal eigenvalue approach. It is shown that two different coupling mechanisms, one involving the symmetric modes and the other involving the anti-symmetric modes, lead to two different instabilities that occur at the same wind speed. The numerical analysis shows that a multi-modal framework and a complete set of aerodynamic coefficients are necessary to correctly estimate the flutter onset. The effect on flutter of structural and aerodynamic characteristics is discussed thoroughly, with a focus on structural modes and on the dependence of the aerodynamic coefficients upon the reduced velocity and upon the angle of attack. [ABSTRACT FROM AUTHOR]

Purpose – The purpose of this paper is to study the structure and dynamic development of a pair of co-rotating trailing vortices, during their formation, interaction and merging, using detailed experimental measurements of the velocity and vorticity fields. Design/methodology/approach – The vortices were generated using two half wings (NACA0030) positioned at equal and opposite angles of attack at the entrance of the test section of an open-circuit, subsonic, wind tunnel. Velocity vector measurements were obtained at Rec = 133,000, on cross-plane grids at several distances from the trailing edges of the wings, using an in-house developed four-sensor hot wire anemometer probe. Findings – The results include cross-plane contour plots of the mean and fluctuating velocity as well as mean vorticity fields. Each of these variables is affected in a different way, providing complementary information on the development of the flow field. After shedding, the two vortices are swept along the stream-wise direction and spiral around each other, thereby developing a braid of two vortices, which then deforms the external flow field. Gradually, the interaction with the external flow field links both vortices together until the final merging and the formation of a new stable linear vortex emerges. Practical implications – Trailing vortices have been rendered particularly important during the past decades, because of increasing traffic density of very heavy aircrafts and several plane “incidents”, which were attributed to the action of the vortex wake. Originality/value – The presented results provide information on the evolution and merging of a pair of vortices formed by a closely spaced differential wing configuration. The vortices interact almost immediately after shedding as expected in flap–flap or flap–wing vortices interaction. [ABSTRACT FROM AUTHOR]

The aerodynamic response to sideslip is crucial for an aircraft's lateral stability, especially for configurations incorporating vortex flow, on which sideslip causes asymmetric vortex behavior and brings extra complexity. In this work, vortex flow over a close-coupled canard configuration under sideslip is simulated using delayed detached-eddy simulation, with the angle of attack varying up to the poststall regime. Sideslip-induced distinctions in both vortex dynamics and canard/wing-vortex interaction are analyzed to explain the nonlinear response of the configuration. The results reveal that the effect of sideslip differs significantly upon different angle-of-attack regimes. At a medium angle of attack, the spanwise extension of the low-pressure area due to windward vortex core expansion results in superior windward lift against the leeward side, but the relatively low pressure retained due to delay of the leeward vortex breakdown is the main motivation that leeward lift surpasses windward at a high angle of attack. The canard vortex trajectory deflected under sideslip induces asymmetric vortex interaction, enhancing the leeward side, which not only leads to an unusual rolling moment toward the higher lift side at a medium angle of attack but also causes an abrupt loss of lateral stability at the poststall angle of attack. A featured demarcation line exists on both the canard and the wing to separate the lift imbalance area, which is found useful in interpreting the nonlinear behavior. [ABSTRACT FROM AUTHOR]

The article provides an overview of an experience with high angle-of-attack (AOA) flight testing at the U.S. Air Force Flight Test Center (AFFTC) in the early 1970s.

Copyright of Electronics & Control Systems is the property of National Aviation University and its content may not be copied or emailed to multiple sites or posted to a listserv without the copyright holder's express written permission. However, users may print, download, or email articles for individual use. This abstract may be abridged. No warranty is given about the accuracy of the copy. Users should refer to the original published version of the material for the full abstract. (Copyright applies to all Abstracts.)

The present study carried out a pure large eddy simulation (LES) on a NACA65 linear compressor cascade (chord-based Reynolds number: 382 000), at an incidence angle of 4°. In this configuration, a corner separation at the juncture of the blade suction side and the end-wall is clearly observed experimentally. For the simulation, 852 CPUs are used in parallel and particular attention is paid to the inflow conditions, coupling LES with the calculation of the inlet boundary layer and associated velocity fluctuations. The shear-improved Smagorinsky model, which has been proved to be competent for turbomachines, is used in this study. Numerical results are carefully compared with pressure measurements, two-dimensional and stereo particle image velocimetry. LES is found to be superior to Reynolds-averaged Navier–Stokes approach in predicting the three-dimensional separated flows in the compressor cascade since pressure coefficients and losses are in very good agreement with the experiment. LES has been proven to simulate also the unsteady behavior of the separation. Finally, the experimental and numerical analysis of the inflow angle reveals high fluctuations of incidence upstream the leading edge of the blade near the end-wall, mostly associated with the incoming wall boundary layer. [ABSTRACT FROM AUTHOR]

Wind tunnel corrections are investigated for two-dimensional low-speed wind tunnel tests that are performed for three similar airfoils for angles of attack ranging from −180° to 180° at Re = 0.75 × 106. Aided by the Blasius theorem, wind tunnel corrections are deduced for the lift, drag and pitching moment of the airfoil at high angles of attack. The wall pressure signature method is applied to determine the strengths of the equivalent singularities. The tunnel wall-induced force and pitching moment are obtained by calculating the force and moment exerted on the equivalent singularities. The maximum correction for drag is determined to be about 50%. The corrected forces and pitching moments for three similar airfoils are coincident with one another. A method to determine an optimum singularities distribution range is presented. The results indicate that the correction method in the paper is effective for airfoil testing at high angles of attack. Copyright © 2014 John Wiley & Sons, Ltd. [ABSTRACT FROM AUTHOR]

A Co-Flow Jet (CFJ) active flow control concept is implemented on the S809 air-foil and numerically investigated by using an in-house code based on Reynolds-averaged Navier-Stokes equations and the Spalart-Allmaras turbulence model, aiming to systematically study the jet effect on the airfoil aerodynamic performance. The solver is validated by comparing the computed results with the baseline experiment measurement. The calculated aerodynamic force and moment coefficients agree fairly well with the experimental data, demonstrating the present solver can predict the attached as well as separated flows around the S809 airfoil with an acceptable precision. The CFJ jet-off geometry of S809 airfoil is simulated to study the jet channel effect on the baseline aerodynamic characteristic, showing that the jet channel could reduce the lift, increase the drag and cause an earlier abrupt stall at angle of attack (AoA) 16.24°. The CFJ jet-on geometry is elaborately studied at three jet momentum coefficient levels, showing that co-flow jet has a significantly positive effect in increasing lift, stall margin, and drag reduction. For the cases with jet momentum coefficients 0.12 and 0.18, the total drag even becomes negative. It is found that the drag from pressure and shear stress of CFJ airfoil is larger than that of baseline, and it is the negative mass-flow-produced drag that significantly reduces the total drag to be below zero. For cases in which AoA are below a critical value (e.g., 20.15° for the case with jet momentum coefficient 0.18), the power required to drive the jet flow decreases when AoA increases, demonstrating that the CFJ concept has an attractive advantage of minimum energy consuming. The present CFJ concept is proved to be a promising and effective active flow control method in the wind turbine application. [ABSTRACT FROM AUTHOR]

Theoretical and experimental investigations to date have assumed that bridge stay cables can be modelled as ideal circular cylinders and that their aerodynamic coefficients are invariant with wind angle-of-attack. On the other hand it has been demonstrated that bridge cables are characterised by local alterations of their inherent surface roughness and shape. Small deviations from ideal circularity result in significant changes in the static drag and lift coefficients with Reynolds number. The present study focuses on the wind-induced response of a full-scale yawed bridge cable section model, for varying Reynolds numbers and wind angles-of-attack. Using passive-dynamic wind tunnel tests, it is shown that the in-plane aerodynamic damping of a bridge cable section, and the overall dynamic response, is strongly affected by changes in the wind angle-of-attack. Using the drag and lift coefficients, determined in static conditions for an identical cable model as the one used for passive-dynamic tests, the in-plane aerodynamic damping is evaluated by employing a one-degree-of-freedom (1 DOF) quasi-steady analytical model. Similarly, it is shown that regions of instability associated with the occurrence of negative aerodynamic damping are strongly dependent on the wind angle-of-attack. [ABSTRACT FROM AUTHOR]

A three-dimensional simulation of hovering flapping wings was performed using an immersed boundary method. This was done to investigate the effects of chordwise wing deformation on three important unsteady aerodynamic mechanisms found in flapping flight, namely Leading Edge Vortex (LEV) shedding, wake capture and clap and fling. A wing was modeled as a flat plate, flapping close to a symmetry plane. Three different deforming chords were defined, a rigid case, a case with maximum deformation at the trailing edge and increased angle of attack (AoA) near the leading edge, and a case with the maximum deformation in the center of the chord and decreased AoA near the leading edge. All cases had zero deformation at the wing root and maximal deformation at the wing tip. A higher AoA near the leading edge resulted in faster LEV buildup and faster buildup of lift. No shedding of the LEV was observed in any of the cases even when deformation caused a high AoA near the leading edge. A distinct dip in lift buildup was observed and shown to be caused by the interaction between the previously shed vortex and the newly developing LEV. This interaction occurred faster when the AoA at the leading edge was increased, and slower when the angle of attack was decreased. Moving the wing closer to the symmetry plane had a positive effect on the cycle average value of CL. This positive effect was reduced however by the earlier interaction between the LEV and the previously shed vortex. [ABSTRACT FROM AUTHOR]

Unsteady aerodynamics imposes additional demands on the modeling of vertical axis wind turbines. Large variations in the angles of attack of the blades cause force oscillations, which can lead to the fatigue-associated problems. Therefore, an accurate estimation of the dynamic loads is essential for the vertical axis wind turbines design. Dynamic stall modeling is in focus because it represents complex phenomena associated with the unsteady flow conditions. The purpose of the study is to find a suitable dynamic stall model for the vertical axis wind turbine conditions. Three versions of the Leishman-Beddoes model are explicitly presented. Additional modifications are implemented for the model to work when the angles of attack change sign and have high amplitudes. All the model parameters are presented. The model is assessed against measured data. The conditions for the simulation tests are close to the vertical axis wind turbine operational conditions. A version of the model, originally designed for low Mach numbers, is the most accurate throughout a number of tests. [ABSTRACT FROM AUTHOR]

Target influence on aerodynamic characteristics in video model deformation experiments has been ignored for high cost of wind tunnel tests, which causes an increase of uncertainty of aerodynamic force data. With the use of computational fluid dynamics, target influence on aerodynamic characteristics of the airfoil RAEE2822 was investigated through three cases to find the influence rules of different parameters. In case I, influence of target thickness ranging from 3H to 9H and target chordwise location ranging from 0.1 to 0.9 on both side surfaces of the airfoil was investigated. In case II, influence of angle of attack ranging from -4° to 8 ° was investigated. In case III, influence of four combination types of targets were investigated with the target thickness ranging from 1H to 9H. Computational results show that the target influence is sensitive to chordwise location and the relation between variation of aerodynamic force and target thickness is nonlinear. Aerodynamic force variation results of different angles of attack also indicate that influence of target on low surface is much smaller than that of target on upper surface when model has a positive angle of attack. [ABSTRACT FROM AUTHOR]

This paper investigates the impact of deploying control surfaces on a model 5 MW offshore wind turbine blade. The wind turbine blade is modeled in 3D based on NREL's 5 MW wind turbine and incorporates varying airfoil shapes and structural twist along the blade length. Lift and aero-dynamic moment at various angle of attack settings are simulated and analyzed. Partial span-wise trailing-edge 3D control surfaces are modeled, simulated and analyzed at various control angles. Results depict the onset of stall around 22 degrees and a larger lift gain versus 10-15 degrees for standard 2D airfoils. Mid-flap controller enhances lift by a factor 2 and mitigates moment a factor of around 7 while end-flap controller enhances lift by a factor of around 1.024 and mitigates moment by a factor of around 2. The results and comparisons obtained serve as one of many instruments for gauging, designing, analyzing, and optimizing wind turbine systems. [ABSTRACT FROM AUTHOR]

The present paper presents oil flow visualisations and pressure measurements over a hemispherecylinder body attached with a forward facing spike at Mach 6 and Reynolds number of 1·38 x 108 at 0° and 5° angle-of-attack. The oil flow pictures depict the separation region in the vicinity of the spike on the hemisphere-cylinder body. The oil flow visualisations will help to locate the reattachment shock wave on the hemisphere-cylinder body and also understand the flow field behavior on the blunt-nosed spike configuration. The pressure measurements over the hemispherecylinder body depend on the shape and the length of the spike. The pressure distributions over the blunt-nosed body show significant influence of the angle-of-attack. The maximum pressure coefficient on the hemisphere-cylinder body is a function of the spike length, shape of the aerodisk and angle-of-attack. The windward and leeward sides pressure variations show dependence of the geometrical parameters of the spike and shape of the spike. The hemisphere and the flat-faced aerodisk cause considerable reduction of pressure leading to decrease of aerodynamic drag compared to the conical spike. [ABSTRACT FROM AUTHOR]

Based on the similarity solution for hypersonic boundary layers, correlative analyses for viscous interaction effect are conducted with numerical method, curves of the wall pressure on the surfaces of the effective body vs the local effective angle of attack and the viscous interactive parameter are presented. The conclusion is that the difference between the wall pressure on the surfaces of the effective body and the real body is proportional to the viscous interactive parameter at bigger local effective angle of attack. Hypersonic flow fields around a lift-body configuration are numerically simulated, the numerical results validated the code and displayed the good agreement with the conclusion of correlative analyses. [ABSTRACT FROM AUTHOR]

Abstract: Using data present in the literature from wind tunnel testing of semi-infinite and infinite rectangular bluff bodies at angle of attack, we develop a simple relation to predict the range of angles of attack where the lift remains approximately constant. This prediction is useful for researchers attempting to develop simplified models of the forces on rectangular bluff bodies. Additionally, prior work has shown that this region of the lift curve is characterized by complex three-dimensional flow, which has consequences for computational simulations. [Copyright &y& Elsevier]

To meet the mission requirement of alternately traversing water-air interface and freely voyaging either in air or water, solve the contradiction between aerodynamic configuration and hydrodynamic configuration, and endure huge water-entry impact, a conformal semi-ring wing configuration is proposed based on the extendable ring wing concept. Basic arrangement of single aero and hydro configuration is framed and the switch method is carried out by the wing's symmetrically rotating around aircraft axis. Aerodynamic and hydrodynamic performance of corresponding configuration is simulated by CFX under designed speed. Numerical simulation results indicate that, when the angle of attack (α) is 0° the lift and drag characteristic of aero configuration is twice as much as the plain wing configuration with the same airfoil, chord length and equivalent wing projective area, when α>12° the lift increasing ratio reduces to ten percent. The performance of hydro configuration is approximate to torpedo configuration, and when the aero configuration voyages in water, the wing drag accounts for 50% of the whole drag, while the wing lift proportion is settled to 50% as |α| > 15° When the absolute value of a is decreasing, lift ratio of the wing to aero configuration increases sharply, at α=3° the ratio reaches 98%. All the results validate the rationality and necessity of the two deployments and morphing scenario of conformal semi-ring wing configuration. [ABSTRACT FROM AUTHOR]

The thrust vectoring control technology is the main technology for the 4th generation fighter. The computational study of interaction flows between vectoring jet and high speed main flow around a simplified fighter configuration is presented. The infections of the parameters such as Mach number of main flow, attack angle, nozzle deflection angle, Mach number, total pressure and total temperature of jet flow are system researched. The physical mechanism of the main flow/vectoring jet flow interaction is also analyzed. Numerical simulation results show that for the simplified fighter configuration, the thrust vectoring jet effect on the high speed main flow is only effects on the aircraft components near the nozzle exit and have little effect on the aerodynamic characteristics of aircraft except at high angles of attack. However, the high speed main flow has strong interaction with thrust vectoring jet flow. With the increasing of the angle of attack and Mach number the decrease of thrust vectoring angle are increasing. [ABSTRACT FROM AUTHOR]

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. A comparison 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. [ABSTRACT FROM AUTHOR]

We present a numerical study on the influence of the Reynolds-number on the lift enhancing effect of a leading edge vortex. Our approach is based on a combination of large-eddy simulations and the immersed boundary technique. We determine the influence of the leading edge vortex on the unsteady lift by simulating a fast pitch-up motion of the plate and studying the lift evolution after holding the flat plate fixed at an angle of attack. Our results suggest that an optimal Reynolds-number exists that maximizes the lift of the leading edge vortex, but that the lift-to-drag ratio is largely independent of the Reynolds-number above a Reynolds-number of Rec > 2000. [ABSTRACT FROM AUTHOR]

In this paper, there are presented some results regarding the mathematical model for helicopter rotor blade aerodynamics, where the sources of unsteady effects were decomposed into perturbations to the local angle of attack and velocity field. The unsteady effects were considered as phase differences between the forcing function and the aerodynamic response, being functions of the reduced frequency, the Mach number, and the mode forcing. For a helicopter rotor, the reduced frequency at any blade element gives useful information about the degree of unsteadiness. The numerical calculus and graphics were made in MAPLE soft medium. [ABSTRACT FROM AUTHOR]

The article deals with numerical and experimental analysis of radio controlled sailplane flight. A Computational Fluid Dynamics programme, Ansys CFX, is used for the determination of aerodynamic forces. Photogrammetry is used for the measurement of speed and glide angle using a video camera. The article presents a comparison of the computed and measured glide ratios. [ABSTRACT FROM AUTHOR]

In this study, we consider various contributions to the forces on an impulsively started finite plate from the perspective of a diagnostic vorticity force theory. The wing plate has an aspect ratio (AR) between 1 and 3, and is placed at low and high angles of attack (α≼ and >20∘), while the Reynolds number is either 100 or 300. The theory enables us to quantify the contributions to the forces exerted on the plate in terms of all of the fluid elements with non-zero vorticity, such as in the tip vortices (TiVs), leading- and trailing-edge vortices (LEV and TEV) as well on the plate surface. This line of force analysis has been pursued for two-dimensional flow in our previous studies. In contrast to the pressure force analysis (PFA), the vorticity force analysis (VFA) reveals new salient features in its applications to three-dimensional flow by examining sectional force contributions along the spanwise direction. In particular, at a large aspect ratio (AR = 3), the force distributions of PFA and VFA show close agreements with each other in the middle sections, while at a lower aspect ratio (AR = 1), the force distribution of PFA is substantially larger than that of VFA in most of the sections. The difference is compensated for by the contributions partly by the edge sections and mainly by the vortices in the outer regions. Further investigation is made fruitful by decomposing the vorticity into the spanwise (longitudinal) component (the only one in two-dimensional flow) and the other two orthogonal (transverse) components. The relative importance of the force contributions credited to the transverse components in the entire flow regions as well as in the two outer regions signifies the three-dimensional nature of the flow over a finite plate. The interplay between the LEV and the TiVs at various time stages is shown to play a key role in distinguishing the force contributions for the plate with a smaller aspect ratio and that with a larger aspect ratio. The present VFA provides a better perspective for flow control by relating the forces directly to the various sources of vorticity (or vortex structures) on or near the wing plate. [ABSTRACT FROM PUBLISHER]

The paper presents a methodology for calculating the unsteady aerodynamic forces characteristics and gust response of the unmanned air vehicle (UAV) model using computational fluid dynamics (CFD) tools, and the efficiency of gust alleviation based on direct force control are also studied. Based on unsteady Euler equations, dynamic responses in vertical gust flow perturbation are investigated. The grid velocity method, which is used to model unsteady flow via grid movement, is introduced to simulate the gust influence. This approach avoids numerical instabilities and decouples the step change in the angle of attack from a pitch rate. And the dynamic chimera grid method is used to simulate the rudder movement of the gust alleviation effect. Numerical results for the NACA0006 airfoil are validated by comparison with analytical results for two dimensional indicial responses. The initial and final lift values of the indicial responses closely match the exact analytical values as given by piston theory and quasi steady thin airfoil theory respectively. The solutions at small times match quite closely the exact solution which is known only for small times given by linear theory. The application of the same method to the UAV model reveals important characteristics of the three-dimensional vertical discrete gust response for the One-minus-cosine gust profile and gusts cause step changes in the angle of attack, and important characteristics of the gust response and resulting flow features are explained. Furthermore, the flow perturbations under dynamic rudder motion are investigated using moving chimera grid technique. Our study shows that with certain designed motions, the flapping rudder can effectively alleviate the freestream gust fluctuation and ensure a stable flight. However, no single parameter can determine whether the flapping rudder can reduce the gust effect. Finally, unsteady aerodynamic loads are computed for the wing of model with the designed rudder motion, the significance of the gust propagation effects is shown. It is found that the direct calculation of the gust response using CFD gives quite accurate results and provides a rich database. In general, this paper shows that the well-designed rudder movement can effectively alleviate the freestream gust fluctuation and ensure high flight qualities. [ABSTRACT FROM AUTHOR]

The growing interest in micro aerial vehicles has brought attention to the need for an improved understanding of the aerodynamics of low-aspect-ratio wings at low Reynolds numbers. In this study, flat plate wings with rectangular and tapered planforms were fabricated with aspect ratios of 0.75, 1, 1.5, and 3, and the aerodynamic loading was measured at Reynolds numbers between 5 x 104 and 1 x 105. Surface tuft visualization was used to observe the interactions between the tip vortices and the leading-edge vortex. The tests were initially conducted at a sideslip angle of 0° and were then repeated for β = 10, 20, and 35° with and without winglets. Measurements made with a six- component force balance showed that a decrease in aspect ratio caused an increase in αstall and Due to image rights restrictions, multiple line equation(s) cannot be graphically displayed. due to the nonlinear lift induced by the interacting flow on the upper wing surface. In addition, the detachment of tip vortices after stall leads to a sudden decrease in drag coefficient as the magnitude of the induced drag drops significantly. At increasing sideslip angles, the effects of the crossflow still contribute to an increase in lift but significantly reduce the pitching moment about the quarter-chord, thus decreasing the wing's ability to recover from angle-of-attack perturbations. These results show that, while the effects of tip vortices and the leading-edge vortex complicate the flowfield around a low-aspect-ratio wing, particularly at increased sideslip angles, their impact tends to improve the aerodynamic performance. [ABSTRACT FROM AUTHOR]