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Research indicates that active control concepts have promise in
mitigating numerous adverse phenomena associated with the aeromechanics
of lifting surfaces. These techniques are being applied to delay stall
of fixed wing aircraft, as well as to eliminate or mitigate vibratory
loads, blade-vortex interaction, and dynamic stall of the flow
about rotorcraft and wind turbine blades. These phenomena are nonlinear
and unsteady for dynamic systems, which add yet another layer of
complexity on the physics of the flow. While a plethora of different
active control techniques is being explored, the use of trailing edge
flaps appears to be one of the more viable and cost-effective concepts.
Static multi-element airfoils and wings have been analyzed
computationally, but little exists on the ability to model these when
the airfoil and flap are dynamic. The costs associated with modeling the
gap between the airfoil and flap have led to approximations where the
flap is modeled only as a morphed tip of the airfoil (no gap). Using a
hybrid Reynolds-Averaged Navier-Stokes/Large-Eddy-Simulation
turbulence technique, an oscillating flapped airfoil has been studied to
determine the influence of modeling the gap on the performance and
acoustic signature of the airfoil. Results are compared with the
experimental data to confirm the validity of the computational approach.
Both attached and separated (dynamic stall) oscillating flows are
examined. The physics within the gap are found to be important for the
airfoil performance when stall is encountered, as well as when acoustic
signatures are required.

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... Using the Chimera technique and a rigid motion to rotate a flap requires considering narrow gaps between the meshes in both flow and span directions, that is, the control surface is "discrete". Although the presence of a gap in flow direction helps to delay stall, the large performance losses at low angles of attack, which usually occur in cruise flight, discourage using discrete flaps [6]. In reality, the gap is usually sealed. ...

... Computations have been performed for Ma ∞ = 0.3, Re ∞ = 30•10 6 and δ = 0°. All simulations have been carried out with the negative version of the Spalart-Allmaras model [16], which is the standard one-equation turbulence model in TAU. ...

... The computational mesh is unstructured and has 62.4•10 6 nodes. As shown in figure 18, it consists in a main mesh block for the wing (drawn in yellow), a mesh block around the tip area (in green) and a mesh block for the split flap which is inside the former one (in pink). ...

The simulation of movable control surfaces is of interest for many applications in aerospace engineering, but it is challenging to perform high-fidelity computations considering them. Generating a new mesh for each deflection angle is computationally expensive, so alternative approaches that allow using only one mesh have been developed in the recent years. This work gives an overview of the available methods for modelling of control surfaces, with special focus on the ones that consider the spanwise gaps between the wing and the control surface. This includes the usage of the Chimera technique combined with mesh deformation as well as a sliding interfaces boundary condition. Their performance has been tested with the DLR MULDICON configuration.

... The configuration of a typical trailing-edge flap is shown in Fig. 17. The use of a fixed flap with a slot for flow control was initially proposed by Page and Glauert in the 1920s, and the primary aim of this technique is to achieve a large lift for an aircraft at a relatively low flying speed [121]. As shown in Fig. 18, a jet-like flow emanating from the gap injects additional turbulence and vortices into the downstream flow over the main airfoil, resulting in the delay or even elimination of the trailingedge vortex. ...

... The airfoil-flap gap for NACA 0012[121].Z.Zhao et al. ...

Floating vertical-axis wind turbines (VAWTs) display considerable advantages over horizontal-axis wind turbines (HAWTs) due to their Omni directionality, better structural scalability, and higher system stability. VAWTs are therefore experiencing a regained interest for use in large-scale offshore wind energy generation. However, the aerodynamic performance of lift-type VAWTs is lower than that of HAWTs. To provide more guidance for the performance improvement of VAWTs, this article reviews the existing approaches to aerodynamic performance enhancement from the perspectives of geometric parameters, flow control methods, blade shape modification, power augmentation devices, hybrid systems, and variable pitch control. Additionally, the findings of various investigations on the performance improvement of VAWTs are summarized.

... As a result, the use of aerodynamic control devices distributed along a wind turbine blade is of great importance in modern wind turbines to control and alleviate the non-uniform loads on the rotor blades 2 . Active control techniques are being applied to delay stall occurrence in aircraft wings, as well as to mitigate vibratory loads and control dynamic stall phenomenon on wind turbine blades 3,4 . 19 (the former for experimental validation and the latter for numerical verification) is considered as the benchmark for pure-pitching motion of an S809 airfoil, so that the airfoil oscillation in this case represents the unsteady motion of the airfoil in the near-stall condition encountering light dynamic stall. ...

... In rotorcraft engineering, the classic plain TEF is implemented which rigidly oscillates around its hinge point. Several experimental and numerical studies have been carried out to investigate the effect of TEFs in rotor blades, helicopter blades in particular4,14,15 . In contrast to the traditional rigid TEFs in rotorcraft engineering, the modern wind turbine blades are equipped with deformable flaps, called Deformable Trailing-Edge Flap (DTEF), attached to the airfoil which have smooth and continuous deformation. ...

Due to the unsteady nature of the flow around horizontal-axis wind turbines, the blades are subjected to severe unsteady and fatigue loads. This necessitates an in-depth aerodynamic analysis of flow control techniques to enhance the performance of a wind turbine as well as the lifetime of its components. Using OpenFOAM package in this study, a series of two-dimensional incompressible simulations are performed to present a deeper insight into the aerodynamic characteristics of an oscillating deformable trailing-edge flap, as a promising flow control device, in a sinusoidal pitching motion of an S809 airfoil. Herein, it is of particular interest to investigate the effects of deformable trailing-edge flap size, oscillation frequency, and the phase shift with reference to airfoil motion on lift and drag hysteresis loops. For this purpose, a pure-pitching motion of an S809 airfoil without flap deflection is considered as the benchmark problem in which the airfoil oscillates in the near-stall region at Re=106. After validation and verification of our simulations through comparison against the corresponding experimental and numerical work, a comprehensive investigation is conducted to study the effects of the aforementioned parameters on the aerodynamic loads. Our results reveal the fact that an out-of-phase deflection of the deformable trailing-edge flap with a frequency equal to the airfoil frequency can significantly mitigate the fatigue load. Under these circumstances, an increase in the deformable trailing-edge flap size can also help the airfoil experience less-severe loads in a cycle of motion. Furthermore, higher values of deformable trailing-edge flap frequency or other values of phase shift except the out-of-phase oscillation cannot alleviate fatigue loads. An airfoil under these conditions can, however, enhance the resultant load required for a blade rotation in the case of low wind periods.

... The (2) Numerical simulations must consider the unique aerodynamic shape of the TEF and model highorder deflection. It is also challenging to simulate the motion between the main blade and the TEF and the flow field near the TEF (such as slot flow [26]) in sufficient detail. Therefore, the authors combined the advantages of the reverse overset assembly technique (ROAT) [27], arbitrary Lagrangian-Eulerian (ALE) numerical simulation, and experience in AFC airfoil simulation in previous research to establish an AFC rotor numerical model and verify the grid independence. ...

Active rotor control of helicopters is the future development direction, and active flap control (AFC) is one of the most promising technologies. However, the numerical simulation of an AFC rotor is challenging. It is necessary to consider the fidelity of the local flow details while dealing with complex shapes and motions. Therefore, few simulations of the flow field and analyses of the influencing parameters have been conducted. In particular, there is a lack of aerodynamic design criteria and recommendations for the AFC rotor. Thus, a new overset assembly algorithm, an arbitrary multilevel moving grid transformation algorithm, and a solver for the unsteady Reynolds-averaged Navier-Stokes equations (URANS) are proposed to establish a suitable numerical method for AFC rotor simulation. The aerodynamic characteristics of the rotor and key influencing factors are systematically analyzed under different flow conditions and design and control parameters, and suggestions for the design of the AFC rotor are provided. The results show that the AFC significantly changes the load distribution of the rotor. The thrust loss of the rotor is approximately 1%, but the offset angle compensates for the loss. The control parameters show relatively consistent trends under different working conditions. The phase is the key control parameter, and the effect on the load is more pronounced when the control frequency is an integral multiple of the rotor’s natural load frequency. Increasing the chord length, span length, and deflection amplitude can also enhance the active control performance.

... Aimed at this problem, an effective solution is proposed by using flaps. Early flaps were widely used in aviation-related fields to increase the lift-to-drag ratio and delay stall at high angles of attack [24,25]. In recent years, flap control as an effective method to reduce blade load and increase blade lift has got a great evolution in the application of VAWT. ...

... They concluded that aerodynamic performances are highly sensitive to the spacing between the cylinder and suction surface. A hybrid Reynolds averaged Navier-Stokes/Large-Eddy-Simulation turbulence technique was adopted by Liggett et al. [38] to study an oscillating flapped airfoil to determine the influence of modeling the gap on the performance and acoustic signature of the airfoil. Their results are compared with the experimental data to confirm the validity of the computational approach. ...

In this paper, we explore the improvement of the aerodynamic characteristics of wind turbine blades under stall conditions using passive flow control with slots. The National Renewable Energy Laboratory (NREL) Phase II rotor, for which detailed simulations and experimental data are available, served as a baseline for assessing the flow control system effects. The position and configuration of the slot used as a flow control system were determined using CFD analysis. The 3D-RANS equations are solved with ANSYS FLUENT using the k-ω SST turbulence closure model. The pressure coefficient for different wind speeds for the baseline configuration is compared to the available experimental data. The comparison shows that CFD results were better for the attached flow. The current work consists of a 3-D CFD modeling of a rotating blade equipped with different flow control systems: single-slot (S-S) and two-slots (T-S). The computation provides a better understanding of the influence of these flow control devices on the performance of wind turbine blades, the control of boundary layer separation, and the rotation effect. These control systems increase the power output by over 60% at high wind speeds with large separated boundary layer regions. For the configuration with the control system, the slot has shown its ability to delay the boundary layer separation. However, the improved aerodynamic performance has been proven for medium and high angles of attack where the flow is generally in the stall condition. The addition of the second slot changed the flow behavior, and an improvement was observed compared to the single slot configuration. The results are helpful for the design and development of a new generation of wind turbine blades.

... In addition, compared with continuous flexible foils, the slot between the main part and the trailing-edge flap of the foil has a strong influence on the downstream flow field. 28 The effect of slots on the flow field is analyzed by utilizing the overset mesh method. For the simulation on turbulent flows, the SST k-x turbulence model is used. ...

The method of oscillating-foil energy extraction can be used to extract kinetic energy from the surrounding flow by a combined pitching and heaving motion of the foil. In order to improve the efficiency of energy extraction, a slotted foil with an active deflecting trailing-edge flap—inspired by the structure of the tail edge of bird wings—is designed. In this study, the unsteady Reynolds-averaged Navier–Stokes equation is solved to investigate the energy extraction performance of an oscillating foil at a Reynolds number of 5.0 × 10⁵. In the numerical simulation, the dynamic overset mesh technology is used in order to ensure the accuracy and convergence of numerical solution. The effect of the deflecting motion of trailing-edge flaps on the efficiency of energy extraction is studied at a range of oscillating frequencies. In addition, the flow control mechanisms of slots on the oscillating foil are revealed by comparing the flow fields of the slotted foil and the NACA0015 foil. The result shows that active deflecting trailing-edge flaps can improve the efficiency of energy extraction over a wide range of oscillation frequencies. The active deflection of trailing-edge flaps increases the energy extraction efficiency of oscillating foils by 21.1% relative to conventional foils under a specific operating condition of oscillating frequency f* = 0.18. A detailed analysis of the flow fields indicates that the slot on the foil can suppress flow separation, while it has a negative effect on the attachment of leading-edge vortices. The deflecting trailing-edge flap enhances the heaving force. Therefore, the energy output and the efficiency of the oscillating foil are enhanced especially at the operating conditions of the oscillating foil without leading-edge vortex shedding.

... A NACA 0012 wing section with a harmonically deflecting TEF was tested in a subsonic wind tunnel by Krzysiak et al. [17] who demonstrated an increase in Cl, max when both the angle of attack of the airfoil and flap deflection angle increase simultaneously. Liggett et al. [22] investigated the impact of an oscillating flap with and without flap gap using a hybrid RANS/LES turbulence model. It was found that the presence of the gaps caused a decrease in aerodynamic performance due to flow recirculation and further confirmed some earlier findings that the oscillating movement drives the unsteadiness in the flow. ...

This work explores the aerodynamic and aeroacoustic responses of an airfoil fitted with a harmonically morphing Trailing Edge Flap (TEF). An unsteady parametrization method adapted for harmonic morphing is introduced, and then coupled with dynamic meshing to drive the morphing process. The turbulence characteristics are calculated using the hybrid Stress Blended Eddy Simulation (SBES) RANS-LES model. The far-field tonal noise is predicted using the Ffowcs-Williams and Hawkings (FW-H) acoustic analogy method with corrections to account for spanwise effects using a correlation length of half the airfoil chord. At various morphing frequencies and amplitudes, the 2D aeroacoustic tonal noise spectra are obtained for a NACA 0012 airfoil at a low angle of attack (AoA = 4°), a Reynolds number of 0.62 × 106, and a Mach number of 0.115, respectively, and the dominant tonal frequencies are predicted correctly. The aerodynamic coefficients of the un-morphed configuration show good agreement with published experimental and 3D LES data. For the harmonically morphing TEF case, results show that it is possible to achieve up to a 3% increase in aerodynamic efficiency (L/D). Furthermore, the morphing slightly shifts the predominant tonal peak to higher frequencies, possibly due to the morphing TEF causing a breakup of large-scale shed vortices into smaller, higher frequency turbulent eddies. It appears that larger morphing amplitudes induce higher sound pressure levels (SPLs), and that all the morphing cases induce the shift in the main tonal peak to a higher frequency, with a maximum 1.5 dB reduction in predicted SPL. The proposed dynamic meshing approach incorporating an SBES model provides a reasonable estimation of the NACA 0012 far-field tonal noise at an affordable computational cost. Thus, it can be used as an efficient numerical tool to predict the emitted far-field tonal noise from a morphing wing at the design stage.

... Much of the recent work focuses either on aerial application of a conventional plain TEF [21][22][23] or on aeroelastic behavior and control strategy of a DTEF. From aerodynamic point of view, Troldborg et al. [24] performed numerical simulations to study the influence of a DTEF on aerodynamic performance of an oscillating RisØ-B1-18 airfoil. ...

Unsteady operating environment of a horizontal axis wind turbine can induce excessive loads on the blades, originating from rapid variations in angle of attack and consequently dynamic stall (DS) occurrence. Therefore, it is of utmost importance to control the flow around a blade by which fatigue damage is likely to happen. Using two-dimensional incompressible unsteady Reynolds-averaged Navier–Stokes equations in OpenFOAM package, a series of simulations are carried out to assess the viability of an oscillating deformable trailing-edge flap (DTEF) in load and DS control on a pitching wind turbine airfoil which experiences deep DS at Re = 420,000. Results reveal whether or not the airfoil is equipped with an oscillating DTEF, DS vortex forms at high angles of attack. The size, strength and traveling of the DS vortex, however, can be influenced by out-of-phase deflection of the DTEF. More effectively, the change in the airfoil camber line during flap oscillation can remarkably affect the pressure distribution around the airfoil, and hence, significant load alleviation and mean lift enhancement are achievable, all of which help the wind turbine performance and enhance the life span of the components. Moreover, a parametric study on flap size and amplitude of deflection together with a comparison between a discrete flap and a DTEF suggests an out-of-phase oscillation of a large gently curved DTEF, up to 30% of the total chord, with similar amplitude and frequency with respect to the airfoil is the best condition under which fatigue load control as well as enhancement in resultant load for a blade rotation can take place.

... The use of high-lift devices extends to the motorsport industry as well, in which they are optimized in-ground proximity for better handling qualities of modern racing cars [1]. More recently, Liggett and Smith [2,3] showed the importance of understanding the flow physics of multi-element airfoils in active flow control applications using unsteady flaps. ...

This work details an experimental investigation on the effects of the variation of flap gap and overlap sizes on the flow field in the wake of a wing-section equipped with a trailing edge Fowler flap. The airfoil was based on the NACA 0014-1.10 40/1.051 profile, and the flap was deployed with 40 deg deflection angle. Two-dimensional (2D) particle image velocimetry (PIV) measurements of the flow field in the vicinity of the main wing trailing edge and the flap region were performed for the optimal flap gap and overlap, as well as for flap gap and overlap increases of 2% and 4% chord beyond optimal, at angles of attack of 0 deg, 10 deg, and 12 deg. For all the configurations investigated, the flow over the flap was found to be fully stalled. At zero angle of attack, increasing the flap gap size was found to have minor effects on the flow field but increased flap overlap resulted in misalignment between the main wing boundary layer (BL) flow and the slot flow that forced the flow in the trailing edge region of the main wing to separate. When the angle of attack was increased to near stall conditions (at angle of attack of 12 deg), increasing the flap gap was found to energize and improve the flow in the trailing edge region of the main wing, whereas increased flap overlap further promoted flow separation on the main wing suction surface possibly steering the wing into stall.

... The flap is hinged at the lower surface of the carbon fiber wing and can move in upward and downward directions over a range of at least 7251. Gap flow between the wing and flap as reported by Liggett and Smith (2013) is prevented by seals covering the gaps on both the suction and the pressure side. The manufacturing tolerance is about 1 mm, meaning that the actual shape differs from the DU96-W-180. ...

Purpose
The purpose of the presented aileron modification analysis is the improvement of the flight handling by eliminating adverse phenomena in the aileron area, such as aileron shaking movements, without the risk of deterioration of flow characteristics during manoeuvres. It was also crucial to reduce aileron forces acting on the control stick.
Design/methodology/approach
Numerical CFD analysis of the aileron system with modifications of sealing in the aileron gap area were performed. The effect of the caulking strip at the upper surface of the aileron gap was determined, as well as caulking at the entrance to the aileron gap on the bottom surface. A solution has also been proposed, consisting of completely closing the aileron gap by using a diaphragm. The three-dimensional flow analysis was carried out, allowing localization of the flow disturbances in the aileron gap at cruising speed. The result of the analysis are the aerodynamic and the hinge moment coefficients determining forces on the control stick, depending on the type of seals.
Findings
It has been shown that the use of subsequent sealing means has a direct impact on the hinge moment value. The results of the CFD analysis showed that the more closed aileron gap is, the higher aileron forces are generated on the control stick. Completely closing the flow in the aileron gap changes the character of the force generated on the control stick.
Practical implications
Through CFD analyses of the aileron gap sealing in the PZL-130 Orlik aircraft, the impact of successive aileron gap sealing on the aileron efficiency was determined. It has been shown that simple change of the aileron gap size by the slat sealing can significantly affect the value of the forces generated on the control stick.
Originality/value
The research using CFD methods allowed to verify the impact of the particular type of aileron gap sealing on the hinge moment value and thus to determine proper sealing configuration for the PZL-130 Orlik aircraft at low computational cost.

Wind turbines are becoming larger to produce more power from the wind in a given area. While the large-size wind turbines are advantageous in terms of generating power, the blades are very heavy and difficult to transport and install. General Electric Co. and the National Renewable Energy Laboratory proposed a new blade design and manufacturing concept that covers the blade with tensioned fabrics. This fabric-covered wind turbine blade is composed of spar-rib structures and covering fabric skins. The present study investigates the aerodynamic effects of fabric skin. A fluid–structure interaction (FSI) analysis was performed about the fabric skin of a large-sized fabric-covered blade. Through static and dynamic FSI analyses, the response of the fabric skin was analyzed according to the wind speeds. The natural frequencies and mode shapes were compared. It was confirmed that the tension of the fabric skin should be increased as much as possible to maintain aerodynamic efficiency, and in addition that the natural frequencies and mode shapes were changed by the wind speeds.

Numerical studies were performed to investigate the mechanism and potential of several active rotors for reducing low-frequency in-plane thickness noise generated by rotating blades. A numerical method coupling the blade element theory, prescribed wake model and Fowcs Williams-Hawkings (FW-H) equation was established for rotor noise prediction. It is indicated that the excitation force on the blade tip can generate anti-noise that to partly cancel the in-plane thickness noise with an appropriate actuation law. Results from the phase, frequency and amplitude sweeps show that the excitation force direction and actuation law are the crucial factors affecting the noise reduction, which determine the noise reduction area in the elevation and azimuth directions, respectively. The active trailing-flap rotor can generate the in-plane excitation force, but because of large lift-drag ratio the anti-noise is mainly from the vertical lift, which is caused by flap deflection similar to a variable camber airfoil. For the harmonic control rotor and active twist rotor, the excitation force is also attributed to the vertical blade lift. The vertical force can reduce the noise near the rotor plane, it will also cause the noise increase in most other areas. Finally, two new active rotors were proposed to generate the in-plane chordwise and spanwise excitation force. With the modified actuation law, the noise in most areas around the rotor was reduced, which improved the acoustic characteristics of rotor significantly.

The present work investigates the feasibility of utilizing the hinge moment of a trailing edge flap (TEF) in monitoring and controlling the unsteady forces on an oscillating airfoil. The aerodynamic forces and hinge moment were measured directly for a NACA 643-618 airfoil model equipped with a TEF in a water channel at a chord Reynolds number (Re) of 46000. At this low Re, the airfoil experiences significant effects of a laminar separation bubble on the suction side. 2D particle image velocimetry measurements were made in the steady and unsteady flow. The results indicated that the TEF hinge moment can be utilized as a localized sensor for the dynamic stall vortex (DSV) shedding into the airfoil’s wake. For this airfoil, controlling the hinge moment provided a promising reduction in the variation in the unsteady normal force despite the significant effort required to control it in the presence of a substantial laminar separation bubble and the development of DSVs.
Graphic Abstract

Rotors with active flap control have considerable potential in reducing vibration and noise and improving aerodynamic performance. However, due to the movement of the flap, there are unavoidable gaps between the components that will lead to significant changes in the aerodynamic characteristics. Moreover, considering the difficulties in motion modeling and the accuracy of the simulation of the flow field in such a narrow gap, carrying out related research is challenging; thus, there has been inadequate targeted research, and that which does exist requires supplementation. To carry out this challenging flow field simulation, the overset assembly algorithm proposed by the author in previous research is adopted in the present study. It is used to successfully assemble the narrow gap, and the accuracy of the simulation is fully verified by comparing the results with the actual experimental results and a grid study. Furthermore, to compensate for the lack of research and experiments on the gap effect, cases considering a complete range of gap widths from an absence of a gap to a width of 10% of the chord length are set up and carried out under the following three case groups: steady cases with a fixed trailing-edge deflection angle, unsteady cases in which only the trailing-edge flap is flapping, and full-motion cases characterized by the periodic flap of the main airfoil and the trailing-edge flap. The results show that the gap increases the drag of the trailing-edge flap and decreases aerodynamic efficiency, especially at low speeds and high angles of attack. Nevertheless, when the gap is unavoidable, there is a range of the gap width that makes unapparent the decrease of aerodynamic efficiency. Moreover, the decrease of aerodynamic efficiency can be reduced as much as possible by a well-designed gap region geometry to ensure that the airfoil and the trailing-edge flap fit together when moving.

In this paper, the numerical simulation was used to investigate the effects of the leading-edge slat installation angles ( β for airfoils from 0° to 40° and β 1 for blades from −20° to 40°) on the aerodynamic characteristics of the airfoil and the wind turbine blade. The chord length of the leading-edge slat is 0.1c (the chord length of the clean airfoil). The horizontal and vertical distances from its center to the leading edge of the clean airfoil are 0.005c and 0.009c, respectively. The results indicated that the lift coefficient could be significantly improved by the leading-edge slat (except β = 40°) when the attack angle exceeded 10.2°. For β = 0°, the lift coefficient increased the most. The trailing vortex of the leading-edge slat played an important role at the process of flow control. It could transfer kinetic energy from the bounder layer to its out-flow region. Furthermore, the vorticities of trailing vortex generated by the leading-edge slat with different installation angles were different, promoting several effects on the airfoil at the different cases. The torque of the blade with leading-edge slat (except β 1 = −20°) was improved significantly as the leading-edge slat trailing-vortices became stronger with the higher wind-speeds.

In order to improve stall characteristics more effectively or eliminate dynamic stall in a revolution of vertical axis wind turbine, three kinds of novel flow-deflecting-gap (FDG) blade were designed and investigated numerically, including two-side FDG blade, toward-outside FDG blade and toward-inside FDG blade. Firstly, the aerodynamic characteristics of isolated flow-deflecting-gap blade were obtained and its effect was illustrated. Then, the investigation on aerodynamic characteristic of novel SB-VAWT was conducted. The obtained results indicated that compared to clean airfoil, FDG enhances the stalling angle of attack (AOA) by 2° and increases lift-drag ratio at high AOA. Furthermore, FDG significantly decreases the frequency of vortex shedding and greatly reduces amplitude of lift oscillation. For a vertical axis wind turbine, toward-inside FDG can efficiently decrease the optimal tip-speed ratio and improve the aerodynamic performance at low tip-speed ratio. Additionally, an increase of 10.21% of torque coefficient is caused by toward-inside FDG. Jet flow in FDG towards leading edge offsets the swirling flow over the suction surface and the trailing-edge vortex produced by bluff edge is also weakened or even eliminated by FDG.

The wind energy has positioned itself as a most promising sustainable energy. The straight-bladed vertical axis wind turbines (SB-VAWTs), as a common turbine for harvesting wind energy, have broad prospects of development. However, the SB-VAWTs are usually influenced by dynamic stall which can cause the aerodynamic losses and fluctuating load. Therefore, the passive flow control (PFC) technique is appreciated for SB-VAWTs due to its low cost and no additional energy consumption. Current paper presents the review of PFC techniques which have been used or are worth being utilized in SB-VAWTs. Furthermore, based on the validation of computational model, a numerical uniform-parameter-criterion study using TSST turbulence model has been conducted to present the research prospects of some novel PFC techniques for SB-VAWTs, including Gurney flap (GF), dimple-GF, leading-edge airfoil-slat (LEAS), flow-deflecting airfoil (FDA), non-circular gap (NCG).

This study is part of a larger effort to create reduced-order models for aerodynamic forces and moments acting on a maneuvering aircraft with moving control surfaces. The aircraft used in this study was inspired by the T-38 jet trainer and includes elevators, rudder, ailerons, and trailing-edge flaps on the wing for landing and takeoff. A hybrid unstructured overset mesh was generated to move these control surfaces and simulate the unsteady flowfields around the aircraft. The static results are first compared to experimental data available at different flap deflection angles, with good agreement obtained at low to moderate angles of attack and deflection angles. Unsteady airloads predictions were then made using the indicial response methods and response functions due to step changes in control surface deflection angles. A time-dependent surrogate model was also used to approximate the response function variation with changes in the angle of attack. The model outputs were then compared with time-accurate simulations of arbitrary control surface motions within the range of data used for model generation. Very good agreement was found between models and computational-fluid-dynamics data for low and high motion rates at low to moderate deflection angles. The results demonstrated that unsteady effects significantly change the amplitude and phase lags of predicted airloads compared with static (or steady-state) predictions.

The validation of fluid–structure interaction solvers is difficult since there is a lack of experimental data. Therefore, in this work an aeroelastic experiment is presented. The focus is on the temporal coupling between fluid and structure dynamics issues in the spatial coupling are eliminated by using a rigid wing. The wing, with a harmonically actuated 0.2c trailing edge flap, has a degree of freedom in the plunge (vertical) direction. The wing has a chord of 0.5 m and is suspended with springs. The wing motion is constrained by a vertical rail system.
For simplicity attached flow is desired and therefore the set angle of attack is α=0°. The Reynolds number is approximately Re=700 000 and the flap deflects over a range of about ±2°. The damped natural frequency of the structure expressed as a reduced frequency is about k=0.194 and measurements are performed for reduced flap frequencies ranging from k=0.1 to k=0.3. Displacements and time dependent aerodynamic forces are measured and for k=0.198 2-D PIV measurements are performed. The planar PIV measurements are used to intrinsically determine the unsteady loads using Noca׳s method.
As expected the aeroelastic problem shows similarities with a viscously damped mass–damper–spring, meaning the maximum excursion of the wing is found near the system eigenfrequency. The lift is dominated by the flap motion and the effective angle of attack due to the motion introduces phase shifts of the lift signal with respect to the flap phase angle.
The experiment has been set up and executed with the necessary precision, but small ambiguities are found in the lift and drag disqualifying the data for validation. Nevertheless the data set provides a clear insight into typical loads and motions and can be used for comparative studies. It can also be used to (re)design future experiments to improve the quality of the data to the desired level of accuracy for validation.

Turbulence and transition modeling still accounts for most of the uncertainties in numerical modeling of complex flows associated with rotorcraft vehicles and components. Computational fluid dynamics (CFD) methods typically cannot capture complex physics with traditional Reynolds-averaged Navier-Stokes (RANS)-based models since majority of physics are transient and occur at different scales. Over the past decade, a resurgence in research related to turbulence modeling has resulted in new large eddy simulation (LES)-based turbulence techniques that have improved computations that involve separated flows. The accuracy of a hybrid RANS-LES technique, first shown to improve turbulent predictions on rotors in the DARPA Quiet Helicopter program, have been increased via locally varying coefficients.

Active trailing-edge flaps are a method of aerodynamic control under extensive research to reduce the detrimental effects of dynamic stall. Physical phenomena are poorly understood in the context of active flaps including vorticity and acoustics, separation, and transition. In addition, discrete trailing-edge flaps create a cavity-like flow within the airfoil-flap gap that can complicate these phenomena. This work has explored the physical response of a static airfoil with a discrete noncontoured oscillating flap over a range of freestream parameters. The effects of attached and separated flows, flap oscillation scheduling, airfoil-flap gap size, and freestream speed have all been investigated. Time-accurate predictions were performed using a hybrid Reynolds-averaged Navier-Stokes/large eddy simulation turbulence model. Trailing-edge stall suppression and an increase between aerodynamic response and deflection input were observed as the flap oscillation frequency increased. The lag between response and input also increased approximately linearly with airfoil-flap gap size. Results indicated the transition was unaffected by the flap oscillations. During the frequency content of flow the unsteadiness was consistent with separated flow driven by the flap. Discrete noncontoured flaps are not recommended; if they are required, the size of the gap should be minimized to maintain performance and reduce lag.

Vortex shedding and the associated noise radiation from a trailing edge were experimentally investigated for a leading-edge slat of a multi-element airfoil at stowed chord Reynolds number Re < 5:9 × 105. A particular focus was on the competition between the instability of the slat boundary layer excited by acoustic feedback and the absolute instability of the wake. Periodic vortex shedding was observed to occur from the slat trailing edge at the Reynolds numbers examined. For Re < 1:9 × 105, the vortex shedding is governed by the absolute instability of the laminar wake of the slat without any distinct tonal noise radiation. For Re > 2:1 × 10 5, however, acoustic feedback becomes pronounced between the trailing-edge noise and the boundary-layer instability waves on the suction surface, so that multiple spectral peaks appear both in the velocity fluctuations and sound pressure. At and around the Reynolds number for the first appearance of tonal noise, Re = 1:9 × 105, both of the instability modes coexist. Beyond Re = 2:1 × 105, the boundary-layer instability waves excited by the acoustic feedback evolve into high-intensity vortices before reaching the trailing edge and suppress the absolute instability of the wake through diminishing the reversed-flow region in the wake. Copyright © 2009 by the American Institute of Aeronautics and Astronautics, Inc.

A study was conducted to demonstrate the ability of a surface-conforming aeroelastic methodology formed from coupling an unsteady Reynolds-averaged Navier-Stokes (URANS) computational fluid dynamics (CFD) solver with a computational structural dynamics (CSD) code to predict pitch/plunge and parabolic camber flutter speeds of a thin symmetrical airfoil in incompressible and compressible flow. The aeroelastic method coupled a finite-element CSD code with a URANS CFD solver at each time step. The fluid-structure interface applied in this effort was developed for three-dimensional morphing rotor blades and modified for application to two-dimensional flexible airfoils. The stability of an airfoil section was evaluated with a single parabolic camber degree of freedom along with pitching and plunging degrees of freedom to verify the methodology.

The present work evaluates the potential of a hybrid RANS-LES method to predict
the unsteady flow over airfoils in static and oscillating motion. The method implemented
(hereafter termed HRLES) blends the k − omega SST RANS model with a localized dynamic
ksgs one-equation LES model (LDKM). The unsteady 2D and 3D flow over a NACA 0015
airfoil is computed to evaluate the model performance. The aerodynamic characteristics
of the static configuration are in reasonable agreement with experimental results. For
the oscillating case, three conditions are simulated: attached flow, mild stall and deep
stall. Two-dimensional simulations are conducted for the three dynamic stall conditions,
and only the deep stall case is simulated in 3D so far. Overall, the unsteady loads for the
attached and mild stall cases show good agreement with experiments. For the mild and the
deep stall cases, the HRLES is able to predict flow separation and vortex shedding during
the downstroke. In general, these results demonstrate the potential of hybrid methods
to correctly simulate complex high Reynolds number flows encountered in aerodynamic
applications.

Helicopter vibration and blade-vortex interaction (BVI) noise are major problems restricting the wider use of helicopters in civil and military applications. Traditional methods based on vibration isolators and absorbers and passive designs of the rotor blade to reduce BVI noise have reached the point of diminishing returns and are increasingly unable to meet the stringent requirements of next generation helicopters. The advent of smart materials such as piezo- ceramics, has opened the possibility of actively twisting the rotor blade using control algorithms in a manner such that new higher harmonic forcing is developed which cancels the existing unsteady higher harmonic aerodynamic forces that are the main sources of vibration and noise on the rotor. Since the main rotor of a helicopter is the principal source of vehicle vibration, active twist control offers possibility of a low vibration helicopter. This paper reviews the literature in active twist rotor control using smart materials.

The paper describes experimental results of controlling flow separation by periodic excitation on the flap of a generic high-lift configuration. The single slotted flap of the two-dimensional test model is equipped with a robust and reliable actuator system that fits inside the flap. The flow is excited using a pulsed wall jet that emanates from the upper surface near the flap's leading edge through a small spanwise-oriented slot. By preventing the flow from separating or by reattaching the separated How, lift and drag are substantially improved, resulting in a lift-to-drag ratio enhancement of 20-25 %. Because of the actuator assembly with spanwise individually addressable segments, the separated flow can be forced to attach only to certain parts of the flap. Local spanwise excitation is thus used to generate a rolling moment without the need to deflect an aileron.

The synergistic use of experiments and numerical simulations can uncover
the underlying physics of airframe noise sources. We focus on the
high-lift noise component associated with a leading-edge slat; flap
side-edge noise is discussed in a companion paper by Streett et al.
(2003). The present paper provides an overview of how slat noise was
split into subcomponents and analyzed with carefully planned
complementary experimental and numerical tests. We consider both tonal
and broadband aspects of slat noise. The predicted far-field noise
spectra are shown to be in good qualitative (and, to lesser extent, good
quantitative agreement) with acoustic array measurements. Although some
questions remain unanswered, the success of current airframe noise
studies provides ample promise that remaining technical issues can be
successfully addressed in the near future.

This paper presents a simulation for high-fidelity aeroelastic analysis of rotating wings with camber-wise structural flexibility and embedded actuators. An unstructured Reynolds-Averaged Navier-Stokes (RANS) computational fluid dynamics (CFD) solver is coupled with a non-linear structural dynamics analysis. The CFD solution uses overset grids to combine the stationary and moving frames of reference. The structural for-mulation expands the conventional one-dimensional beam representation with additional degrees-of-freedom to capture plate-like cross-sectional deformations while allowing an arbitrary distribution of active and pas-sive materials in the cross section. Motion and forces on the non-coincident fluid and structural grids are transferred using a finite-element-based interpolation, along with a least-squares fit for extrapolations. Trim and convergence to periodic response are assisted by a low-order analysis that is also discussed. Finally, as an initial verification of the implementations, results from the low-order and CFD-based solutions are compared for a rigid-airfoil rotor in forward flight.

noise has proven to be a useful guide for elucidation of the physics of flow-induced noise generation over the last five years. This process, relying on a close interplay between experiment and computation, is described and demonstrated here on the archetypal problem of flap-edge noise. Some detailed results from both experiment and computation are shown to illustrate the process, and a description of the multi-source physics seen in this problem is conjectured.

Effects of trailing-edge flap gaps on rotor performance are investigated using a high fidelity, coupled computational fluid dynamics (CFD) - computational structural dynamics (CSD) analysis. Both integral flap (the flap is an integral part of the blade such that there are no physical gaps at the flap ends) and discrete flap (the flap is a separate entity with physical gaps in the spanwise and chordwise directions) are examined on an UH-60A rotor at high speed forward flight condition. A novel grid deformation scheme based on the Delaunay graph mapping is developed and implemented to allow the CFD modeling of the gaps with minimal distortion of mesh around the flap gap regions. This method offers an alternative to the traditional approach of modeling such configurations using overset meshes. The simulation results show that the effectiveness of the flap is minimally lost with spanwise gaps - the penalty on rotor performance is of the order of 1% compared to the integral flap. On the other hand the chordwise gaps significantly degrade the benefits of active flap on rotor performance due to the flow penetration between the upper and lower surfaces of the flap.

Stereoscopic PIV technique is applied to a low-speed flow around a simple wing-flap configuration in a large-scale wind tunnel to demonstrate the applicability and capability of the technique to flow fields around high-lift systems. Also, CFD analyses by a Reynolds-averaged Navier-Stokes code are conducted to evaluate usefulness of PIV in the CFD code validation. Test results suggest usefulness of the stereoscopic PIV for investigation of complex, three-dimensional flow fields around a high-lift device configuration, including tip vortices while two-dimensional PIV might be more suitable for detailed measurements of near-wall flow such as slot and cove flows. Comparisons between the PIV and CFD results show that the stereoscopic PIV is a useful tool for validation of CFD code applied to high-lift device design while improvements in terms of measurement accuracy and capability of measuring near-wall velocity are required. Especially, focusing on flows featuring slot and wake flows of the flap, which are considered sensitive to CFD solver and turbulence model, is appropriate in critical CFD validations. Based on the velocity field measurements for different model configurations, it is inferred that the flow field around a flap is sensitive to both flap slot gap and flap deflection angle in the parameter ranges tested in this study, and therefore, large gap height or deflection angle easily results in a large-scale separation with a highly turbulent recirculation zone on the flap upper surface.

The flow around the ONERA RA16SC1 three-element airfoil was numerically investigated using three different computational strategies: the Unsteady Reynolds-Averaged Navier-Stokes Equations (URANS), the Detached Eddy Simulation (DES) and the Implicit Large Eddy Simulation (ILES). Two different numerical schemes were employed: the Jameson's Central Difference Scheme (CDS) and the third-order weighted essentially non-oscillatory (WENO) scheme in conjunction with the HLLC Riemann solver. The numerical results were compared with experimental measurements consisting of pressure plots and PIV data. Similar to the experiment, ILES predicts the flow separation in the flap region whereas URANS and DES do not. In the slat cove, the DES and ILES provide similar flow results, including the Reynolds stresses. Copyright © 2009 by the American Institute of Aeronautics and Astronautics, Inc.

This paper presents numerical simulations of the flow around a NACA 0015 airfoil at static and dynamic stall. The treatment of these configurations is a very challenging task for CFD applications. The turbulent flow around the static and in pitch oscillation airfoil is computed using different approaches: 2D RANS, 3D RANS and DES methodologies and with finer and finer meshes in order to try to reach a space converged solution. The main conclusion of the paper is that the prediction of static and all the more dynamic stall is not mature with present modeling capabilities.

The ability to predict SMART active trailing edge flap rotor loads is explored in this study. Full-scale wind tunnel data recently acquired in the NASA Ames 40-by 80-Foot Wind Tunnel are compared with analytical results from CAMRAD II. For the 5-bladed rotor, two high-speed forward flight cases are considered, namely, a 0 deg flap deflection case and a 5P, 2 deg flap deflection case. Overall, the correlation is reasonable, with the following exceptions: the torsion moment frequency and the chordwise bending moment are underpredicted. In general, the effect of the 5P, 2 deg flap motion is captured by the analysis, though there is overprediction in the neighborhood of the 105 deg and 120 deg azimuthal locations. Changes to the flexbeam torsion stiffness are also briefly considered in this study, as this stiffness will be updated in the future. Finally, the indication is that compressibility effects are important, and this suggests that computational fluid dynamics might improve the current correlation.

To address the complex multidisciplinary nature of rotorcraft analysis, high-fidelity computational fluid and structural dynamics models have been developed and coupled for an advanced technology active rotor. Significant advancements have been made in both modeling disciplines to allow for complex bearingless flapped rotors. Comparisons are made between CFD/CSD and comprehensive (lifting-line, free-wake) analyses and experimental data for the Boeing SMART rotor. Flap phase sweeps for 0, 2, 3, 4, and 5/rev flap inputs are investigated in relation to the zero flap deflection baseline at a nominal cruise condition. Changes in performance, aerodynamic and structural loads, control power, and noise are studied. Details of the high-fidelity flowfield solution, including flap gap effects and wake visualizations, are also presented.

The vortex shedding caused by compressible subsonic flow along a wall cavity has been investigated using a Large-Eddy Simulation (LES)-based turbulence modeling technique that is embedded within a legacy Reynolds-Averaged Navier-Stokes (RANS) solver to assess the improvement in the prediction of the flow field and acoustic of cavity flows beyond the application of classic RANS turbulence models. Numerical simulations applying two-equation Kinetic-Energy Simulation (KES), sub-grid scale hybrid-RANS LES (HRLES-sgs), and Menter k - w shear-stress transport (SST) turbulence methods have been carried out and compared with experiment and LES results. Important frequencies of the flow are determined, illustrating the abilities of advanced turbulence modeling to improve these predictions when compared to RANS models. Evaluation of the influence of the grid, time step and simulation period shows the sensitivity of the predictions to these parameters.

The vortex shedding generated by compressible subsonic flow interacting with a wall cavity has been investigated using large-eddy-simulation-based turbulence techniques embedded within a legacy Reynolds-averaged Navier-Stokes solver. Cavity simulations using hybrid turbulence approaches seek the accuracy of large-eddy simulation by providing filtering and modeling of subgrid-scale turbulence with the cost of traditional Reynolds-averaged Navier-Stokes. Simulations applying differing techniques of hybridization of the Menter k-omega shear stress transport Reynolds-averaged Navier-Stokes approach include detached eddy simulation (DES-SST), blended subgrid-scale turbulence models (GT-HRLES), and a self-adjusting large-eddy-simulation very-large-eddy-simulation technique (KES) provide an understanding of differing hybrid approaches. Cavity flow results from Reynolds-averaged Navier-Stokes and hybrid simulations are compared with experiment and large-eddy simulation predictions. Evaluation of important flow characteristics illustrates the abilities of these advanced turbulence modeling techniques compared with traditional Reynolds-averaged Navier-Stokes models. Examination of the influence of the grid, time step, and simulation period demonstrates the sensitivity of the aerodynamic and aeroacoustic predictions to these parameters. In particular the subgrid-scale blended model, GT-HRLES, shows significant improvement in the ability to capture the acoustic signatures and flowfield features on a Reynolds-averaged Navier-Stokes or very-large-eddy-simulation grid compared with the other models.

Current rotorcraft research to increase flight speed or to alleviate adverse physical phenomena expand the Mach/angle-of-attack envelope in which the rotor blades operate. For example, rotor blades will experience large areas over the rotor disk where reverse-flow effects cannot be neglected during the design and analysis of an efficient rotor at high advance ratios. A cost-effective alternative to extensive experimental analyses is the use of computational fluid dynamics codes to quantify the behavior of airfoils at high and reverse angles of attack, as well as to add to the knowledge of the behavior of airfoils when they are immersed in these flows. Numerical experiments have been performed with correlation to experimental databases that examine the ability of computational fluid dynamics to accurately model airfoil characteristics at these angles of attack. It is observed that the use of recently developed hybrid Reynolds-averaged Navier-Stokes and large-eddy simulation turbulence methods result in a significant improvement in the ability of computational fluid dynamics to predict the characteristics of airfoils in these angle-of-attack regimes. Modeling of the airfoil trailing edge is more sensitive when reverse-flow angles of attack are considered.

The effects of a harmonically deflected trailing-edge flap, actuated at different start times and amplitudes but with frequency different from the airfoil motion, on the aerodynamic loads of an oscillating NACA 0015 airfoil were investigated experimentally at Re=2.51×105. Both in-phase and 180° out-of-phase flap deflections, relative to the airfoil motion, were tested. The results show that there was a large change in the hysteretic behavior of the dynamic load loops, and that the formation and detachment of the leading-edge vortex (LEV) were not affected by the flap motion, while the low pressure signature of the vortex was affected by the flap actuation start time. The later the flap actuation the larger the change in the strength of the LEV. The present flap control scheme was also found to be as effective as that achieved by a pulsed ramp flap motion, but with a reduced number of control parameters.

A preliminary study of the control of the dynamic-load loops of an oscillating NACA 0015 airfoil through an 18%c LEF (leading-edge flap) and a 25%c TEF (trailing-edge flap), deflected dynamically and independently, was conducted. Both upward and downward deflections actuated at 1 st = Oμ were tested to maximize the effects of the flap motion on the behavior of the dynamic-stall vortex (DSV). The downward LEF motion suppressed leading edge separation and eliminated the occurrence of the DSV, leading to a minor reduction in C i, maxbut a considerably improved poststall lift condition, compared with the baseline airfoil. On the other hand, the upward LEF motion caused an earlier formation and shedding of the DSV and, subsequently, an increased C i,max and negative peak C m value. The downward TEF deflection always increased the lift. The formation and detachment of the DSV was, however, largely unaffected.

In this paper an extensive parametric study concerning the effect of flap and slat riggings on the 2-D high-lift flow past a three-element airfoil system is presented. The numerical approach for solving the Reynolds-averaged Navier-Stokes equations uses an implicit finite volume scheme of second order accuracy in space on a patched multizonal grid. The Spalart-Allmaras one-equation turbulence model is employed. Six design parameters have been investigated comprising the deviation, gap, and overhang of the slat and of the flap whose settings are centered about the values in practice for takeoff. Good agreement with experiments is obtained for prestall angles of attack. Computations show that both C-l and C-l/C-d have an optimum with every design parameter. The trends in the high-lift flow observed are in accordance with both experiments and computations reported in the literature. C-l and C-l/C-d are found to be more sensitive to deviations and gaps than to overhangs.

Unsteady aerodynamic loads on NACA 0012 airfoil with a trailing-edge flap were measured in wind tunnel and calculated from a simple theoretical model. The airfoil model of 0.18 m chord length used in wind-tunnel test was oscillating in pitch about an axis located at 35% chord length from the airfoil leading edge. The length of trailing-edge flap was 22.6% of airfoil chord. The Hap was also deflecting harmonically, but with frequency different from airfoil pitching motion. The influence of phase delay between airfoil angle of incidence and flap deflection at the beginning of the motion was considered. The theoretical method used for calculation of unsteady airfoil loads is based on two-dimensional, inviscid, incompressible flow model at subsonic Mach numbers. The expressions for unsteady aerodynamic loads calculations on the airfoil and on the flap were obtained in a closed form using distribution of flow velocity potential along the airfoil chord and along the flap length. Lift and aerodynamic moment measured in the wind tunnel were compared with results of calculations. The correlation between experimental and theoretical results is adequate.

A multielement airfoil designed for helicopter application has been tested for compressible dynamic stall behavior and has been proven to he a robust dynamic stall-free concept. This slotted airfoil has operated into poststall areas without the dynamic stall vortex that is normally present whenever airfoils are tested beyond their static stall boundary. Point diffraction interferogram images of the dynamic flow over the airfoil are presented, showing details of the flow development during the oscillation cycle, and instantaneous pressure distributions on the airfoil and slat during dynamic airfoil motion are included.

Vortex shedding and the associated noise radiation from a trailing edge were experimentally investigated for a leading-edge slat of a multi-element airfoil at stowed chord Reynolds number Re < 5:9 × 105. A particular focus was on the competition between the instability of the slat boundary layer excited by acoustic feedback and the absolute instability of the wake. Periodic vortex shedding was observed to occur from the slat trailing edge at the Reynolds numbers examined. For Re < 1:9 × 105, the vortex shedding is governed by the absolute instability of the laminar wake of the slat without any distinct tonal noise radiation. For Re > 2:1 × 10 5, however, acoustic feedback becomes pronounced between the trailing-edge noise and the boundary-layer instability waves on the suction surface, so that multiple spectral peaks appear both in the velocity fluctuations and sound pressure. At and around the Reynolds number for the first appearance of tonal noise, Re = 1:9 × 105, both of the instability modes coexist. Beyond Re = 2:1 × 105, the boundary-layer instability waves excited by the acoustic feedback evolve into high-intensity vortices before reaching the trailing edge and suppress the absolute instability of the wake through diminishing the reversed-flow region in the wake. Copyright © 2009 by the American Institute of Aeronautics and Astronautics, Inc.

A zonal hybrid Reynolds-averaged Navier-Stokes large-eddy simulation (RANS/LES) approach, called zonal-DES, used to handle a two-dimensional high-lift configuration with deployed slat and flap is presented. This method allows to reduce significantly the cost of an accurate numerical prediction of the unsteady flow around wings compared to a complete LES. Some issues concerning grid generation as well as the use of zonal-detached-eddy simulation for a multi-element airfoil are discussed. The basic planar grid has 250,000 points and the finest spanwise grid has 31 points with Delta z/c = 0.002. The effort is geared toward detailed comparison of the numerical results with the Europiv2 experimental particle image velocimetry data including both mean and fluctuating properties of the velocity field (Arnott, A., Neitzke, K. P., Agocs, J., Sammer, G., Schneider, G., and Schroeder, A., "Detailed Characterisation Using PIV of the Flow Around an Aerofoil in High Lift Configuration," EUROPIV2 Workshop on Particle Image Velocimetry, Springer, Berlin, 2003). The results also provide an insight into the real unsteady nature of the flow around a three-element airfoil that cannot be reproduced by classical RANS models. The current calculation displays extremely complex flow dynamics in the slat and flap coves like the ejection process through the gaps oiseveral vortices issued from the impingement of the free shear layers on the lower walls of the different elements.

The control of cavity flows has been investigated by the means of Large Eddy Simulations. The computations have been carried out on unstructured meshes to assess the efficiency of two passive acoustic oscillation suppression devices: the rod-in-crossflow and the flat-top spoiler. Despite a sustained interest and many experiments, a clear explanation for observed reduction in the flow-induced structure load is still missing. This work explores different hypotheses: the modification of the mean field and its linear stability properties, a pure deflection effect of the separated shear layer, or scale coupling between the rod wake and the turbulent mixing layer over the cavity. The aim here is to enhance the experimental database and provide leads towards a better understanding of the phenomena. The selected test-case is a cavity of length/depth ratio equal to 5, at Mach and Reynolds number of M∞=0.85 and ReL=7.106, respectively.

Stall hysteresis discovered in the wind-tunnel performance of a GA(W)-2 airfoil with a 25% chord slotted flap is examined further by using the data obtained for lift, pitching moment, surface pressure distribution, and the hot-film velocity vector. Test cases include 30- and 40-deg flap deflections, each having an optimum and narrow gap at a chord Reynolds number of 2.2 x 10(6) and a Mach number of 0.13. The flap optimized to produce the highest C-Imax for each flap angle apparently did not have a proper contour and nose location for the slot flow to function effectively at off-design conditions. It is shown that suction pressures over the flap, suppressed by a thickening wing wake at stall, are not reversible to their prestall values within the decreasing alpha side of loop. It is suggested that the flap design include the use of a new flap parameter called the slot flow angle to describe the slot flow orientation and a pressure recovery factor to select a proper contour for the flap upper surface.

A investigation has been made in the N.A.C.A. 7-by-10 foot wind tunnel of a large-chord N.A.C.A. 23012 airfoil with several arrangements of venetian-blind flaps to determine the aerodynamic section characteristics as affected by the over-all flap chord, the chords of the slats used to form the flap, the slat spacing, the number of slats, and the position of the flap with respect to the wing. Complete section data are given in the form of graphs for all the combinations tested. The optimum arrangement of the venetian-blind flap was a combination in which the flap was located near the wing trailing edge. These arrangements of the venetian-blind flap were superior to any flaps previously tested for producing lift and giving low drag coefficients at high lift coefficients. The wing with this flap, however, had very large pitching- moment coefficients. When operated as split flaps, the venetian-blind flaps were inferior to the simple split flap in producing lift.

The Tactical Technology Office of the Defense Advanced Research Projects Agency (DARPA) initiated the Helicopter Quieting Program (HQP) in 2004 to develop high fidelity, state-of-the-art computational tools for designing advanced helicopter rotors with reduced acoustics perceptibility and enhanced performance. A critical step towards this achievement is the development of high-end rotorcraft prediction codes capable of assessing a wide range of helicopter configurations and operations for future rotorcraft designs. This includes novel next-generation rotor systems that incorporate innovative passive and/or active elements to meet future challenging military performance and survivability goals. Phase I of the HQP program involved development of prediction methodologies ("tools") by coupling computational structural dynamics (CSD) to computational fluid dynamics (CFD) modeling codes. Participants (vendors) included joint efforts by Stanford University/University of Maryland (SM), as well as Georgia Institute of Technology/Pennsylvania State University (GP) and Teledyne Sciences Corporation (TSC). Phase I was primarily geared towards validating these prediction tools for conventional rotors currently in use by the fleet. Results from the vendors demonstrated significant improvements in prediction accuracy and correlations [1] over classical comprehensive methods in all aspects of the aerodynamics, structural and acoustics responses of the rotor. Phase Ib1 was initiated in 2007 to demonstrate the robustness of HQP tools in the characteristics of unconventional rotor designs that utilize innovative on-blade active controls for dynamic tuning. An active flap rotor currently under-development at Boeing (i.e. the Boeing SMART rotor) was selected as the candidate for this code validation effort.

Results of a study using a passive approach to recover the loss of lift that occurred when a variable droop leading edge (VDLE) airfoil was used to successfully control compressible dynamic stall by attaching a small Gurney flap to its trailing edge are reported. Gurney flaps of different heights were tested. The airfoil performance was evaluated by measuring the unsteady pressures while it executed a sinusoidal pitch-up maneuver over a range of Mach numbers from 0.2 to 0.4, at different reduced frequencies, with both static and dynamic leading edge droops. Not only was the “lost lift recovered completely with a 1% chord-height Gurney flap, the drag and moment coefficients were also dramatically reduced and a lift-to-drag ratio greater than 10 was achieved, making it an acceptable choice for this purpose. The improved performance is explained through the basic fluid mechanics of the problem by discussing the various pressure distributions and the surface vorticity fluxes derived from these.

The ability to model accurately and efficiently unsteady aerodynamic effects for actively controlled trailing-edge flaps (ACFs) is crucial for practical application of such systems for vibration and noise reduction as well as performance enhancement. Two-dimensional unsteady airloads due to oscillating flap motion are calculated and compared using various computational fluid dynamics (CFD) codes and a CFD-based reduced-order model (ROM). This ROM is based on the rational function approximations approach, which yields a state-space, time-domain aerodynamic model suitable for incorporation into comprehensive rotorcraft simulation codes. The accuracy of this model is demonstrated across a practical range of unsteady flow conditions encountered by active flaps. Two Reynolds-averaged Navier-Stokes solvers (CFD++ and OVERFLOW) are employed in conjunction with various turbulence models, including large eddy simulation based models, so as to examine code independence. Flow physics associated with three-dimensional effects, flap hinge gap, as well as compressibility effects are also examined.

Turbulent flows over lifting surfaces exhibiting trailing-edge vortex shedding often cause adverse and complex phenomena, such as self-induced vibration and noise. In this paper, a numerical study on flow past a blunt-edged two-dimensional NACA 0015 section and the same section with various base cavity shapes and sizes at high Reynolds numbers has been performed using the unsteady Reynolds-averaged Navier–Stokes (URANS) approach with the realisable κ–ε turbulence model. The equations are solved using the control volume method of second-order accuracy in both spatial and time domains. The assessment of the application of URANS for periodic trailing-edge flow has shown that reasonable agreement is achieved for both the time-averaged and fluctuating parameters of interest, although some differences exist in the prediction of the near-wake streamwise velocity fluctuation magnitudes. The predicted Strouhal numbers of flows past the squared-off blunt configuration with varying degrees of bluntness agree well with published experimental measurements. It is found that the intensity of the vortex strengths at the trailing-edge is amplified when the degree of bluntness is increased, leading to an increase in the mean square pressure fluctuations. The numerical prediction shows that the presence of the base cavity at the trailing-edge does not change the inherent Strouhal number of the 2D section examined. However, it does have an apparent effect on the wake structure, local pressure fluctuations and the lift force fluctuations. It is observed that the size of the cavity has more influence on the periodic trailing-edge flow than its shape does.

Measured, open loop and closed loop data from the SMART rotor test in the NASA Ames 40-by 80-Foot Wind Tunnel are compared with CAMRAD II calculations. One open loop high-speed case and four closed loop cases are considered. The closed loop cases include three high-speed cases and one low-speed case. Two of these high-speed cases include a 2 deg flap deflection at 5P case and a test maximum-airspeed case. This study follows a recent, open loop correlation effort that used a simple correction factor for the airfoil pitching moment Mach number. Compared to the earlier effort, the current open loop study considers more fundamental corrections based on advancing blade aerodynamic conditions. The airfoil tables themselves have been studied. Selected modifications to the HH-06 section flap airfoil pitching moment table are implemented. For the closed loop condition, the effect of the flap actuator is modeled by increased flap hinge stiffness. Overall, the open loop correlation is reasonable, thus confirming the basic correctness of the current semi-empirical modifications; the closed loop correlation is also reasonable considering that the current flap model is a first generation model. Detailed correlation results are given in the paper. Notation c m Pitching moment coefficient C T Helicopter thrust coefficient KTEF Flap hinge stiffness, ft-lb/rad M Mach number NP Integer (N) multiple of rotor speed Per rev Per revolution RmPtn NASA wind tunnel Run "m" Point "n" α Angle of attack α s Rotor shaft angle µ Rotor advance ratio σ Rotor solidity ratio Sign Convention Chordwise moment, + tip toward trailing edge. Flap deflection, + trailing edge down. Flap lift, + up; flap chordwise force, + toward leading edge. Flatwise moment, + tip up. Pitch link load, + in tension. Torsion moment, + leading edge up.

Modern wind turbines are steadily increasing in size, with recent models boasting rotor diameters greater than 120 m. Wind turbines are subjected to significant and rapid fluctuating loads, which arise from a variety of sources including turbulence, tower shadow, wind shear and yawed flow. Reducing the loads experienced by the rotor blades can lower the cost of energy of wind turbines. ‘Smart rotor control’ concepts have emerged as a solution to reduce fatigue loads on wind turbines. In this approach, aerodynamic load control devices are distributed along the span of the blade, and through a combination of sensing, control and actuation, these devices dynamically control the blade loads.
This research investigates the load reduction capabilities of smart rotor control devices, namely trailing edge flaps (TEFs), in the operation of a 5 MW wind turbine. A feedback control approach is implemented for load reduction, which utilizes a multiblade coordinate transformation. Single input–single output control techniques are employed to determine the appropriate response of the TEFs based on the blade loads. The use of TEFs and this control approach is shown to effectively reduce the fatigue loads on the blades, relative to a baseline controller. The load reduction potential is also compared to an alternative individual pitch control (IPC) approach, in the time and frequency domain. The effects on the pitch and power systems are briefly evaluated, and the limitations of the analysis are assessed. Finally, a combined approach that uses both TEFs and IPC is evaluated. Copyright

Detailed measurements on mean flow and turbulence around a multielement airfoil model have been made using pressure and hot-wire probes. The results obtained in two test cases at chord Reynolds number of 3 x 10 exponent 6 and freestream Mach number of 0.2 show a number of features of the complex flows that are important for accurately modeling these flows by numerical methods. Many parts of the shear flow deviated vastly from classical flows, and the interaction with the external flow is very strong.

Measurements and simulations are presented of the flow past a tailplane research airfoil which is designed to show a mixed leading-edge trailing-edge stall behaviour. The numerical simulations were carried out with two flow solvers that introduce transition prediction based on linear stability theory to RANS simulations for cases involving laminar separation bubbles. One of the methods computes transition locations across laminar separation bubbles whereas the other assumes transition onset where laminar separations occur. For validation of the numerical methods an extensive measurement campaign has been carried out. It is shown, that the methodology mentioned first can simulate the size of laminar separation bubbles for angles of attack up to where the separation bubble and the turbulent separation at the trailing edge are well behaved and steady in the mean. With trailing edge separation involved, the success of the new numerical procedure relies on the diligent choice of a turbulence model. Cases with large 3D flow structures inside the turbulent trailing edge separations in windtunnel experiments for high angles of attack are compared and analysed along with 2D and 3D steady RANS calculations that model the measurement section of the windtunnel.

The self-excitation mechanism of the acoustic diametral modes of an axisymmetric internal cavity–duct system is studied for a Mach number range up to 0.4. The effect of cavity dimensions on the excitation mechanism is investigated experimentally and numerically. Experiments are conducted on three cavity depths and six cavity lengths for each depth. Numerical simulations of the mode shapes are also performed to determine the effect of cavity dimensions on the particle velocity field of the diametral modes. For all the tested configurations, the diametral modes are strongly excited at relatively low Mach numbers (as low as 0.1). The pulsation amplitude at resonance is found to increase as the cavity becomes shorter or deeper, relative to the main pipe diameter. The test results provide new insights into the excitation mechanism of diametral modes, the effect of the cavity length to depth ratio on the Strouhal numbers of acoustic resonances caused by various shear-layer modes of the cavity, and into the effect of the particle velocity field of the acoustic modes on the mode selectivity mechanism which determines the dominant acoustic mode during resonance.

An experimental study was made of the flow regimes associated with the cove regions of a multielement airfoil. These regions occur when the leading edge slats and trailing edge flaps on the wings are extended to generate high lift during takeoff and landing of the aircraft. The results include mean velocity and turbulence intensity data, measured between the slat and wing and between the wing and flap, together with comprehensive pressure distributions around the three airfoil components. This information should be of value in the development of numerical models for predicting the flow around high-lift airfoils.

Experiments have been performed at low speeds documenting certain broad aspects of hysteresis on a mildly
swept wing under high-lift conditions. Aerodynamic load measurements were carried out at two values of incidence,
10.5 and 15.5 deg, and flap deflections of 20 and 30 deg, essentially under quasi-steady conditions. Two-dimensional
particle image velocimetry (PIV) was utilized to measure the mean velocity vector field in the rear part of the wing
and flap. These velocity measurements have revealed the complex nature of separated flows on the wing and flap.
Also, in general, the slot flow angle determined by geometric considerations is very different from those actually
inferred from PIV measurements.

Several issues relating to the application of Chimera overlapped grids to complex geometries and flowfields are discussed. These include the addition of geometric components with different grid topologies, gridding for intersecting pieces of geometry, and turbulence modeling in grid overlap regions. Sample results are presented for transonic flow about the Space Shuttle launch vehicle. Comparisons with wind tunnel and flight measured pressures are shown.

Analysis of Development of Dynamic Stall based on Oscillating Airfoil Experiments

- L Carr
- K Mcalister
- W Mccroskey

Carr, L., McAlister, K., McCroskey, W., 1977. Analysis of Development of Dynamic Stall based on Oscillating Airfoil Experiments. Technical Report NASA TN D-8382. NASA.

Systematic Investigations of the Effects of Planform and Gap Between the Fixed Surface and Control Surface on Simple Flapped Wings

- B Gothert
- C Rober

Gothert, B., Rober, C., 1949. Systematic Investigations of the Effects of Planform and Gap Between the Fixed Surface and Control Surface on Simple Flapped Wings. Technical Report TM-1206. NACA.

Boeing-Smart Test Report for DARPA Helicopter Quieting Program Unsteady airfoil with a harmonically deflected trailing-edge flap

- B Lau
- N Obriecht
- T Gasow
- B Hagerty
- K Cheng
- B Sim

Lau, B., Obriecht, N., Gasow, T., Hagerty, B., Cheng, K., Sim, B., 2010. Boeing-Smart Test Report for DARPA Helicopter Quieting Program. Technical Report TM 2010-216404. NASA. Lee, T., Su, Y., 2011. Unsteady airfoil with a harmonically deflected trailing-edge flap. Journal of Fluids and Structures 27, 1411–1424.

Active Flap Control of the Smart Rotor for Vibration Reduction

- S R Hall
- R V Anand
- F K Straub
- B H Lau

Hall, S.R., Anand, R.V., Straub, F.K., Lau, B.H., 2011. Active Flap Control of the Smart Rotor for Vibration Reduction. Technical Report ARC-E-DAA-TN590. NASA.

non-contoured gaps are not recommended for design due to large associated performance losses at low angles of attack Flow-excited resonance of trapped modes of ducted shallow cavities

- K Ziada

Although gap flows can delay stall, non-contoured gaps are not recommended for design due to large associated performance losses at low angles of attack. References Aly, K., Ziada, S., 2010. Flow-excited resonance of trapped modes of ducted shallow cavities. Journal of Fluids and Structures 26, 92–120.

Zonal hybrid RANS-LES method for static and oscillating airfoils and wings Flow regimes in the cove regions between a slat and wing and between a wing and flap of a multi-element airfoil

- M Sanchez-Rocha
- M Kirtas
- S Menon

Sanchez-Rocha, M., Kirtas, M., Menon, S., 2006. Zonal hybrid RANS-LES method for static and oscillating airfoils and wings. In: 44th AIAA Aerospace Sciences Meeting and Exhibit, Reno, NV. Savory, E., Toy, N., Tahouri, B., Dalley, S., 1992. Flow regimes in the cove regions between a slat and wing and between a wing and flap of a multi-element airfoil. Experimental Thermal and Fluid Science 5, 307–316.

Application of the chimera overlapped grid scheme to simulation of space shuttle ascent flows

- P Buning
- S Parks
- W Cham
- K Renze

In search of physics: the interplay of experiment and computation in airframe noise research, part 2

- M Khorrami
- D Lockard
- B Singer
- M Choudhari
- C Streett