Delay Flow Separation Research Articles

Experimental Investigation of Low Reynolds Number Flow Around a Serrated NACA 0015 Airfoil

The study of low Reynolds number flows has triggered the interest of many researchers. This concern is attributed to the presence of various vortex structures that significantly affect aerodynamic performance. The study of such phenomena offers an opportunity to better understand and predict the behavior of fluid flows in three dimensions, leading to improved design strategies for aerodynamic systems. This study intends to examine the performance of different NACA 0015 airfoil configurations namely, baseline, serration, comb, comb-serration and poro-serrated, emphasizing on their influence on aerodynamic characteristics and flow structure. Experimental work is performed to provide valuable insights into the behavior of the flow field and the underlying physical phenomena. The current study has shown that the mean streamwise velocity of the serrated, combed, and comb-serrated configurations exhibit more stable flow patterns compared to the baseline airfoil. These patterns suggest a promising approach for delaying flow separation. In contrast, the poro-serrated model exhibited a more disordered flow pattern. Remarkably, when subjected to a 10-degree angle of attack, all the modified trailing-edge designs showcased minimal separation zones compared to the baseline configuration. At the same angle, the baseline's shear layer disintegrates, while the poro-serrated model displays increasing disturbance and separation. Conversely, the comb-serrated and serration models exhibit a smaller increase in unsteadiness. Furthermore, the study indicated that Musou black paint exhibits a higher rate of light absorption compared to flat black paint, but it demonstrates higher reflection levels on a poro-serrated surface. Collectively, the outcomes of this investigation hold significant implications for enhancing aerodynamic system designs.

Improvement of turbulent flow in a bend using rotational guide vane: A shape optimization

The present study aims to minimize velocity disruption using rotational guide vane (RGV) before the 90° pipe bend where the curvature ratio is r/D = 1.0 at Re = 3 × 104. The combination of computational fluid dynamics and genetic algorithm is used as a tool for the geometric optimization of RGV. The SST k–ω turbulence model was preferred in flow analysis. The number of blade, the blade angle, and the location of RGV are the geometric parameters to be optimized. A new evaluation method of the velocity distribution after the bend is proposed. The objective function is defined as velocity distribution mean squared error that is based on the comparison of the velocity distribution on the cross section after the bend with the fully developed flow in a straight pipe under the same conditions. As a result of a series of optimization processes, velocity distribution at the 90° bend exit is compared between optimized RGV and without guide vane. The results show that the optimized RGV improves velocity distribution at the bend exit. Compared to the without guide vane case, the optimized RGV has delayed flow separation from α = 37° to 56°, and the flow reattachment point moves from x/D = 0.5 to 0 at the bend exit. Thanks to optimized RGV, the counter-rotating Dean vortices merged to form a single vortex at the center of the pipe. In addition, significant reduction in turbulent kinetic energy was observed, approximately 50% when using the optimized RGV compared to the without guide vane case.

Control effect of superhydrophobic grooves on flow-induced noise generated by flow around cylindrical shell at large Reynolds number

Flow-induced noise is an important factor affecting the quiet performance of underwater vehicles. Superhydrophobic surfaces are an emerging technology for underwater vehicles. In this study, a superhydrophobic surface is innovatively applied to the flow-induced noise control of underwater cylindrical shells. Alternating no-slip and no-shear surfaces are used to simulate the superhydrophobic surface with spanwise superhydrophobic grooves so that the flow regime and flow-induced noise of a no-slip cylinder are compared with the superhydrophobic cylinder under a high Reynolds number. The results show that the superhydrophobic surface can effectively delay flow separation and control the size of the wake shedding vortex. The flow-shedding vortices mainly affect the flow-induced noise in the lower frequency range, which is consistent with the vortex shedding frequency. The radiation characteristics of the flow-induced noise generated by the fluctuation pressure are mainly influenced by the eigenfrequency of the model in the range of 100 Hz–5000 Hz. Moreover, the superhydrophobic surface can effectively reduce the flow-induced noise and change its radiation directivity at both high and low frequencies by controlling vortex shedding and reducing the fluctuation pressure, respectively. The findings reported here shed new light on the flow-induced noise control of underwater vehicles.

Experimental Investigations on Flow Control of the Rotor via the Synthetic Jets in Forward Flight

To study the effects of synthetic jet control on the aerodynamic performance of a rotor in forward flight, we conducted a series of experiments with varying rotor rotation speeds and free stream velocities. In the test, we used a six-component balance and a PIV system and designed a blade with a particular structure that covered the frame. The experimental results revealed that the synthetic jet could effectively delay flow separation over the blade and enhance the aerodynamic efficiency of the rotor. Moreover, we investigated how different jet parameters influenced the flow control effects of synthetic jets on the rotor’s aerodynamic characteristics. We drew some valuable conclusions from our analysis. In forward flight, the jet located closer to the leading edge of the blade had a stronger impact on improving the aerodynamic performance of the rotor. The jet with a 90° jet angle increased the rotor normal force by 225%, which was the maximum possible increase, while the jet with a 30° inclined angle had the best control effects on preventing flow separation in the retreating blade. Our study provides valuable insights into the use of synthetic jets for rotor flow control and suggests possible applications for improving rotorcraft performance and stability.

Alula Effect on Aerodynamic Performance of Small UAVs Wing Configuration

The application range of small UAVs for civil purposes such as aerial imaging, agricultural pest control and, corps management as well as drone delivery has been increasing throughout recent years. Promising results are observed from recent studies which support the effectiveness of the bio-inspired approach in improving the aerodynamic efficiency of small UAVs. However, very limited studies have investigated the effect of Alula on the aerodynamic performance of wing geometries. This research investigates the effects of employing Alula on the aerodynamic efficiency of the wing to overcome the challenge of small UAVs efficiency degradation. Numerical and experimental investigations are conducted on 6 different Alula configurations. The findings of the investigations shows a reduction in drag for wings and an increase in aerodynamic efficiency by 9% for wings with Alula compared to the conventional wing. It is observed that using Alula configuration helps in delaying flow separation by adding momentum into the flow which resulted in total drag reduction without having a significant effect on the lift. Consequently, the aerodynamic efficiency is increased, and this can be utilized in drone industry to increase the endurance of the flight.

Phononic-subsurface flow stabilization by subwavelength locally resonant metamaterials

The interactions between a solid surface and a fluid flow underlie dynamical processes relevant to air, sea, and land vehicle performance and numerous other technologies. Key among these processes are unstable flow disturbances that contribute to fundamental transformations in the flow field. Precise control of these disturbances is possible by introducing a phononic subsurface (PSub). This comprises locally attaching a finite phononic structure nominally perpendicular to an elastic surface exposed to the flowing fluid. This structure experiences ongoing excitation by an unstable flow mode, or more than one mode, traveling in conjunction with the mean flow. The excitation generates small deformations at the surface that trigger elastic wave propagation within the structure, traveling away from the flow and reflecting at the end of the structure to return to the fluid-structure interface and back into the flow. By targeted tuning of the unit-cell and finite-structure characteristics of the PSub, the returning waves may be devised to resonate and reenter the flow out of phase, leading to significant destructive interference of the continuously incoming flow waves near the surface and subsequently to their attenuation over the spatial extent of the control region. This entire mechanism is passive, responsive, and engineered offline without needing coupled fluid-structure simulations; only the flow instability’s frequency, wavelength, and overall modal characteristics must be known. Disturbance stabilization in a wall-bounded transitional flow leads to delay in laminar-to-turbulent transition and reduction in skin-friction drag. Destabilization is also possible by alternatively designing the PSub to induce constructive interference, which is beneficial for delaying flow separation and enhancing chemical mixing and combustion. In this paper, we present a PSub in the form of a locally resonant elastic metamaterial, designed to operate in the elastic subwavelength regime and hence being significantly shorter in length compared to a phononic-crystal-based PSub. This is enabled by utilizing a sub-hybridization resonance. Using direct numerical simulations of channel flows, both types of PSubs are investigated, and their controlled spatial and energetic influence on the wall-bounded flow behavior is demonstrated and analyzed. We show that the PSub’s effect is spatially localized as intended, with a rapidly diminishing streamwise influence away from its location in the subsurface.

Leading-Edge Active Flow Control Enabled by Curved Fluidic Oscillators

A new design of fluidic oscillator, in which the oscillator is curved along its primary flow direction, allows for a sweeping jet to be placed near the leading edge of a wing as an active flow control device. Previous studies of fluidic oscillator applications have shown their promise as an active flow control device by increasing the lift over a wing and preventing flow separation. However, the planar extent of the flat fluidic oscillators creates challenges when trying to optimally implement them in a confined geometry. The application of these new curved oscillators is experimentally investigated on the leading edge of an upswept NACA0018 airfoil full-span wing. Global load cell measurements of lift, drag, and pitching moment are presented and analyzed using suction surface static pressure taps, and particle image velocimetry data are presented. The effects of active flow control on lift, drag, and stall at a Reynolds number of 100,000 are documented for a variety of oscillator chordwise locations near the leading edge and compared to traditional vortex-generating jets. While further aft-positioned oscillators produced a larger increase in maximum lift, more forward locations resulted in better poststall performance and delayed flow separation over the wing.

Influence of free-stream turbulence intensity on static and dynamic stall of a NACA 0018 aerofoil

In many engineering applications, aerofoils experience elevated free-stream turbulence levels. The present work experimentally investigates the effect of free-stream turbulence on the aerodynamic characteristics of a sinusoidally pitching NACA 0018 aerofoil at the transitional Reynolds number of 2.8×105. Wind tunnel tests are conducted in quasi-static and dynamic conditions at different turbulence intensities between 0.3% and 8.2%, considering reduced frequencies up to 0.1. The dynamic experiments investigate multiple angle-of-attack ranges in order to quantify the influence of free-stream turbulence in the attached-flow regime, in the stalled regime, and in-and-out of stall. The study demonstrates that high free-stream turbulence drastically changes the flow physics over the aerofoil for both static and dynamic conditions, producing large deviations in the lift and moment coefficients. The quasi-static experiment performed at low free-stream turbulence features a large stall hysteresis linked to the breakdown of the leading-edge laminar separation bubble, whereas static hysteresis is not present at higher turbulence levels. Moreover, dynamic experiments at high angles of attack show a strong dependency on the incoming turbulence intensity, which is found to delay flow separation during the upstroke and enhance reattachment during the downstroke. The work also reveals the intrinsic difficulty of predicting the dynamic stall behaviour under different turbulence conditions, and gives insight into why the existing empirical dynamic stall models are unlikely to succeed in accurately predicting the aerodynamic loads across different onset turbulence intensities.

Entropy analysis and parametric optimization on the Nano fluid flow and heat transfer of a pin-fin heatsink with the splitter using the two-phase mixture model

The nanofluid (NF) flow rate and concentration (φ) in a pin-fin heat sink with the splitter (WS) were numerically investigated and the optimum Re and φ values were determined by applying a parametric optimization analysis on the thermal and frictional entropy generation rates (S˙th and S˙fr). To this end, the discretization of the governing equation was done by the second-order upwind scheme, the finite volume method. The results were compared with those obtained for the pin-fin heatsink without the splitter (WOS). Based on the findings, the intensification of Re and φ reduces S˙th by almost 43-45% and 10-13%, respectively, for both WS and WOS heatsinks. An insignificant difference between S˙th of two configurations was obtained (0.37-2.93%). Although S˙fr diminishes by 12% as φ intensifies from 0 to 1%, the escalation of Re from 500 to 2000 intensifies S˙fr by 2312% and 4618% in WS and WOS heatsinks, respectively. In addition, S˙fr of WS heatsink is nearly 43-48% lower than that WOS heatsink due to using splitters. The absence of the splitter in the WOS heatsink leads to a delay in flow separation from the pin-fin bodies and deteriorates the heat dissipation, thereby increasing the CPU temperature and S˙fr intensification.

Wall-Resolved Large-Eddy Simulation of Near-Stall Airfoil Flow at

A wall-resolved large-eddy simulation (LES) of the near-stall flow around the Aerospatiale A-airfoil at is conducted. The present LES shows typical Reynolds-number effects compared to the previous LES at , such as the increase in lift coefficient, decrease in boundary-layer thickness, delay of turbulent flow separation, and upstream shift of transition location. Among these Reynolds-number effects, the difference in the development of momentum displacement thickness is focused on and further analyzed based on the integral relation of the boundary layer. This analysis reveals that the different mechanisms of the laminar–turbulent transition between the two Reynolds-number cases cause an initial difference in the momentum displacement thickness. This initial difference is then amplified downstream by the deceleration effects of the mean flow. These results suggest that proper estimation of the laminar and transitional flows near the leading edge is crucial for predicting the trailing-edge stall phenomena and their dependency on the Reynolds number.

Two-Dimensional URANS Numerical Investigation of Critical Parameters on a Pitch Oscillating VAWT Airfoil under Dynamic Stall

In this paper, a thorough 2D unsteady computational fluid dynamic analysis was performed on a pitching airfoil to properly comprehend the dynamic stall and aerodynamic forces. The computational software ANSYS Fluent was used to solve the unsteady Reynolds-averaged Navier–Stokes equations. Low Reynolds number flows were modeled using the k-ω shear stress transport turbulence model. Aerodynamic forces, fluid flow structures, and flow separation delay angles were explored as a function of the Reynolds number, reduced frequency, oscillation amplitude, and mean angle of attack. The maximum aerodynamic forces, including lift, drag, and the onset of the dynamic stall, were all influenced by these variables. The critical parameters that influenced the optimum aerodynamic forces and ended up causing dynamic stall delay were oscillation amplitude and mean angle of attack. The stall angle was raised by 9° and 6°, respectively, and a large increment in the lift coefficient was also noted in both cases. Additionally, for the highest Reynolds number, a considerable rise in the maximum lift coefficient of 20% and a 28% drop in drag coefficient were observed, with a 1.5° delay in the stall angle. Furthermore, a significant increase of 33% in the lift force was seen with a rise of 4.5° in the stall angle in the case of reduced frequency.

Development and interaction of vortices over a very low aspect-ratio wing under pitch-up motion

The vortical structures over a thin rectangular wing with a very low aspect ratio of 0.277 were investigated in a wind tunnel at an effective Reynolds number of $3 \times 10^6$ . When applying pitch-up motion pivoted at mid-chord, the maximum lift angle was increased with an increase in the pitch rate, but the maximum lift coefficient was reduced. The pitching motion also caused delay in the vortical development over the wing, which was increased with an increase in the pitch rate. The delay in the leading-edge vortex development due to the pitching motion was nearly identical to that in the tip vortex development, indicating that the dynamics of the leading-edge vortex was strongly influenced by the tip vortex. This was confirmed by particle image velocimetry measurements, which demonstrated that the tip vortex over a very low aspect-ratio wing induced strong downwash to influence the development of the leading-edge vortex during the pitching motion, which led to a delay in flow separation.

Aerodynamics of a three-dimensional bionic morphing flap

Avian wings are characterized by intricate feather movements to manipulate airflow and enhance aerodynamic performance. The wing trailing-edge undergoes continuous gradual transformations to enact camber reflexes. A three-dimensional bionic flap inspired by the avian wings is proposed. The flap undergoes spanwise morphing to generate seamless wave-like deformations at the trailing-edge. The performance of S809 airfoil fitted with various spanwise morphing flaps, and conventional flaps, is studied. Comprehensive investigations are performed to analyze the aerodynamic forces, pressure-fields, and velocity evolution across the airfoils fitted with different flaps, subjected to various deflection angles and angles of attack, at the Reynolds number of 106. The analyses exhibit significance of variable mean airfoil-camber on the aerodynamic performance. Comparative analyses demonstrate the superiority of morphing flaps over the conventional ones with regards to- lift increment, drag reduction, and enhanced glide ratio. Further, flow-field visualization and surface-flow measurements reveal unique 3D-evolution of the inflow in streamwise and spanwise directions. The spanwise morphing significantly delays flow separation and aids turbulence suppression. The downstream wake-velocity profiles are also found to be considerably reformed by the spanwise morphed flap topology. Overall, the study establishes supremacy of spanwise morphing over the conventional design in terms of achieving optimal aerodynamic performance.

Effect of turbulence intensity on aerodynamic characteristics of extended trailing edge airfoil

PurposeThis paper aims to study the extended trailing edge airfoil for a range of angle of attack at different intensities of turbulence.Design/methodology/approachIn this paper, an experimental study on NACA 0020 airfoil with thin extended trailing edge modification of amplitude of h = 0.1c, 0.2c and 0.3c at the Reynolds number of 2.14 × 105 are tested. The research was carried out for an angle of attack ranging from 0° = α = 35° for the turbulence intensity of 0.3%, 3%, 5%, 7% and 12%. From the experimental readings, the surface pressures are scanned using a Scanivalve (MPS2464) pressure scanner for a sampling frequency of 700 Hz. The scanned pressures are converted to aerodynamic force coefficient and the results are combined and discussed.FindingsThe airfoil with the extended trailing edge will convert the adverse pressure gradient to a plateau pressure zone, indicating the delayed flow separation. The CL value at higher turbulence intensity (TI = 12%) for the extended trailing edge over perform the base airfoil at the post-stall region. The maintenance of flow stability is observed from the spectral graph.Practical implicationsA thin elongated trailing edge attached to the conventional airfoil serves as a flow control device by delaying the stall and improving the lift characteristics. Additionally, extending the airfoil's trailing edge helps to manage the performance of the airfoil even at a high level of turbulence.Originality/valueDistinct from existing studies, the presented results reveals how the extended trailing edge attached to the airfoil performs in the turbulence zone ranging from 0.3% to 12% of TI. The displayed pressure distribution explains the need for increasing trailing edge amplitude (h) and its impact on flow behaviour. The observation is that extended trailing edge airfoil bears to maintain the performance even at higher turbulence region.

The effect of trip wire on transition of boundary layer on a cylinder

The effect of height of a trip and its location on the transition of boundary layer on a cylinder is studied using large eddy simulations for 2×103≤Re≤5×105. The Reynolds number, Re, is based on the free stream speed and diameter of the cylinder (D). Two modes of transition are observed: (a) natural, for a relatively small trip of height dT/D=0.25%, via formation of a laminar separation bubble (LSB) and (b) direct, for a large trip of height dT/D=1.0%, wherein the formation of LSB is bypassed and the trip disturbs the flow enough to cause separation of the boundary layer and its subsequent turbulent reattachment. Transition delays the final separation leading to a very significant reduction in drag, often referred to as drag crisis. The delay is more for natural as compared to direct transition. Consequently, the drag at the end of crisis is lower for natural transition. The 1.0% trip at 78° leads to a more delayed flow separation than one at 55° from the front stagnation point. The drag crisis takes place in two stages for a cylinder with trip. During each of the two stages, asymmetric transition on the two sides results in generation of circulation and lift force. The effect of trip is felt even by the non-trip side. The cylinder experiences a relatively large “reverse lift” during the second stage of drag crisis. While natural transition is accompanied by intermittency of LSB, direct transition is associated with intermittency in laminar vs turbulent attachment of the flow following its separation at the trip.

Numerical Assessment of a BAY-Type Model for different Vortex Generator Shapes Applied on a Wind Turbine Airfoil

Modern wind turbines are increasingly equipped with vortex generators in the inner part of the rotor blade. On large blades it is usual to install hundreds of these small passive devices to delay flow separation. Therefore, affordable numerical methods to include vortex generators in the design process are required. One very promising method based on the addition of source terms in the flow equations is the BAY model. In this work a BAY-type model is presented and assessed for three different vortex generator shapes (rectangular, delta and cropped-delta) placed on a DU 97-W-300 airfoil and computed with steady-state Reynolds-averaged Navier-Stokes numerical methods. In order to quantify the circulation of the main streamwise vortex, a method to identify the vortex core based on the λ2-criterion is presented and shows good results. The computations reveal very good agreement with the experimental data for attached flow in terms of lift and pressure distribution while drag is inherently underestimated by the BAY model due to the inviscid modelling of the vortex generators. It is displayed that this underestimation is proportional to the vortex generator surface area. Concerning the different shapes, it is shown that the delta shape reveals larger deviations from the fully resolved case than the other geometries and therefore requires higher mesh resolution.

Delay or control of flow separation for enhanced aerodynamic performance using an effective morphed surface

Delay or control of flow separation for enhanced aerodynamic performance using an effective morphed surface

Spanwise-Discontinuous Grooves for Separation Delay and Drag Reduction of Bodies with Vortex Shedding

Suitably shaped grooves, placed transverse to the flow, can delay flow separation over curved surfaces. When grooves are fully extruded in the spanwise direction, they may reduce the drag of boat-tailed bodies with vortex shedding, but with the drawback of increasing the spanwise correlation of the vortex shedding. We investigate herein the effect of spanwise-discontinuous grooves through Large Eddy Simulations. A systematic analysis is carried out on the effect of the number, N, of grooves that are present for N equally long portions of the total spanwise length of the boat-tail. Discontinuous grooves further reduce the drag compared with the full-spanwise-extruded groove. Increasing N produces an improvement of the flow-control-device performance, whose maximum value is reached for N=3, corresponding to a spanwise extension of the groove roughly equal to the body crossflow dimension. Above this value, no further improvements are found. The maximum drag reduction is equal to 25.7% of the drag of the boat-tail without grooves and to 17.7% of the one of the boat-tail with the full-spanwise-extruded groove. The lowest drag value occurs for the least correlated vortex-shedding in the spanwise direction. The reduction in the correlation is mainly related to a flow separation line that is less regular in the spanwise direction.

Experimental study of the effect of a slat on the aerodynamic performance of a thick base airfoil

Abstract. There is continuous effort to try and improve the aerodynamic performance of wind turbine blades. This experimental study focusses on the addition of a passive slat on a thick airfoil typically used in the inboard part of commercial wind turbine blades. Nine different slat configurations are considered, with both a clean and tripped main airfoil. The results are compared with the performance of the airfoil without slat, as well as the airfoil equipped with vortex generators. It is found that, when the airfoil is clean, the increase in lift-to-drag ratio due to the presence of a slat is larger than when vortex generators are used. This is also true for the tripped airfoil but only at small angles of attack. As expected, in all configurations, the presence of the slat delays flow separation and stall. Finally, for a clean airfoil and small angles of attack, the slat decreases the lift-to-drag ratio of the main airfoil only. By contrast, as the angle of attack increases, it seems that the slat changes the flow field around the main airfoil in such a way that its lift-to-drag ratio becomes larger than for the airfoil without slat. These effects are less pronounced when the airfoil is tripped. This work helps to better understand the role of a slat in improving the aerodynamics of blade sections. It can also be used to validate simulation tools in the field.

Theoretical breakthrough in the dynamics of a jet in a free-stream flow around a corner

For the first time, the laminar jet in a free-stream flow around a corner with/without a moving wall condition (as it assists the flow) is studied theoretically. The corner may contract the streamlines or expand them depending on the angle of inclination. In the context of incompressibility, it is presented a composite non-similar transformation of the Navier-Stokes (N-S) equations in Cartesian space to approximately decipher the nature of the boundary layer flow over the entire physical domain; i.e. 0 ≤ x, y < ∞. The transformed Partial Differential Equations (PDEs) are then solved by an in-house MATLAB code, which employs an implicit algorithm of a tridiagonal form with quasi-linearization technique. For a stationary wall, as the negative angle of the corner meets Falkner-Skan (F-S) limitation and expands beyond, the predictive model starts to show a critical location, which reveals directly the flow separation point. This critical location moves toward the jet origin as the angle of inclination becomes more negative. Here, these locations are extracted as a function of the inclination parameter in order to manifest the suitability of a jet to delay boundary layer separation. A similar scenario also applies to the moving wall case. In this case, it was recorded that a moving wall condition delays separation for a considerable distance. It should be pointed out that up to date, there is no theoretical investigation on this specific flow geometry despite the many associated applications; and the limited available reports on an analogous flow geometry are those from the Computational Fluid Dynamics (CFD) and on a case by case basis. More specifically, the present geometry is quite similar to the flow passing the way down to the trailing edge of an airfoil, with a jet generator mechanism such as Dielectric Barrier Discharge (DBD) plasma actuator to delay flow separation on the backside of the airfoil.Here, by the use of the theoretical non-similarity concept empowered by some recent advancements (see Jafarimoghaddam, 2020 and Jafarimoghaddam, 2021), the fundamental problem of a jet discharged in the lower boundary of a freestream flow over a wedge is successfully solved.This theoretical breakthrough suggests promising avenues for the future design of airfoils with moving parts/flaps as an innovative step forward to enhance maneuverability of the flying objects.