Airfoil Flow Separation Research Articles

On the Generalization Capability of a Data-Driven Turbulence Model by Field Inversion and Machine Learning

This paper discusses the generalizability of a data-augmented turbulence model with a focus on the field inversion and machine learning approach. It is highlighted that the augmented model based on two-dimensional (2D) separated airfoil flows gives poor predictive capability for a different class of separated flows (NASA wall-mounted hump) compared to the baseline model due to extrapolation. We demonstrate a sensor-based approach to localize the data-driven model correction to tackle this generalizability issue. Furthermore, the applicability of the augmented model to a more complex aeronautical three-dimensional case, the NASA Common Research Model configuration, is studied. Observations on the pressure coefficient predictions and the model correction field suggest that the present 2D-based augmentation is to some extent applicable to a three-dimensional aircraft flow.

Sinusoidally Pitching Foils in Transonic Flow: Insights into Flutter from Time Accurate Simulations

The impact of shock formation and vortex dynamics on airfoil flutter in the transonic regime is investigated through a series of two-dimensional time-accurate simulations of a sinusoidally pitching airfoil at a Reynolds number of 10,000. To gain insight into transonic flutter in the fully nonlinear regime, an energy-based approach that employs prescribed kinematics is utilized. Flutter stability boundaries are identified as a function of the Mach number and the reduced velocity, the latter of which is a surrogate for structural stiffness or flight speed. The results show a subcritical instability near Mach numbers of M∞=0.7 for a range of oscillation frequencies. Additionally, three shock-induced mechanisms are identified that influence the transfer of energy from the flow to the airfoil. The primary mechanism is flow separation triggered by the generation of a lambda shock. The influence of the lambda-shock dynamics on airfoil flow separation is determined for different Mach numbers, providing insight into the well-known “transonic dip” observed in flutter boundaries.

Experimental study on Q-learning control of airfoil trailing-edge flow separation using plasma synthetic jets

In this paper, a continuous Q-learning algorithm is deployed to optimize the control strategy of a trailing-edge airfoil flow separation at a chord-based Reynolds number of 2×105. With plasma synthetic jets issued at the middle chord and a hot wire placed in the separated shear layer acting as the actuator and sensor, respectively, a high-speed reinforcement learning control at an interaction frequency of 500 Hz is realized by a field-programmable gate array. The results show that in the Q-learning control, the controller only needs several seconds to elevate the instantaneous reward to a level close to the final mean reward, and convergence of the control law typically takes less than 100 s. Although the relative drag reduction achieved by Q-learning control (10.2%) is only slightly higher than the best open-loop periodical control at F∗=4 (9.6%), the maximum power saving ratio is improved noticeably by 62.5%. Physically, Q-learning control creates more turbulent fluctuations, earning more rewards by increasing the transition possibilities toward high-value states. With increasing penalty strength of plasma actuation, the final control laws obtained from Q-learning exhibit a decreasing number of active states. Detailed comparisons between the open-loop and Q-learning control strategies show that the statistics of the controlled velocity fields remain similar, yet the turbulent fluctuations contributed by the vortex shedding mode are reduced by constant-frequency plasma actuation.

Wind-tunnel tests of synthetic jet control effects on airfoil flow separation

To find out the comprehensive principles of synthetic jet control effects on airfoil stall characteristics, fundamental contrastive wind-tunnel tests were conducted systematically. By using six-component balance, Particle Image Velocimetry (PIV) technology and boundary layer probe, the measurements of model aerodynamic forces, the whole velocity field over airfoil and velocity profiles in the boundary layer were conducted, respectively. Based upon the experimentally parametric analyses of synthetic jet control on airfoil, it is concluded that the incline angle of synthetic jet has an important impact on both the flow in boundary layer and aerodynamic forces of airfoil. A tangential jet inject energy and accelerate the velocity of flow in inner boundary layer thus has a better control effects on delay stall of airfoil when the momentum coefficient of jet is relatively large, and the normal jet helps to enlarge the thickness of boundary layer which is proved to have better control effects on flow separation control and improving the aerodynamic characteristics of airfoil when the momentum coefficient of a jet is relatively small.

Experimental study on airfoil flow separation control via an air-supplement plasma synthetic jet

An air-supplement plasma synthetic jet (PSJ) actuator increases the air supplemental volume in the recovery stage and improves the jet energy by attaching a check valve to the chamber of a conventional actuator. To explore the flow control effect and mechanism of the air-supplement actuator, via particle image velocimetry experiments in a low-speed wind tunnel, the flow field and boundary layer characteristics of a two-dimensional airfoil surface under different actuation states were compared for different attack angles and jet orifices. The experimental results show that, compared with the conventional actuation state, the jet energy of the air-supplement PSJ is higher and the indirect mixing effect of the counter-vortex sequence produced by the jet-mainstream interaction is stronger. Furthermore, the boundary layer mixing effect is better, which can further suppress flow separation and improve the critical flow separation attack angle. Moreover, increasing the jet momentum coefficient can enhance the flow control effect. The findings of this study could provide guidance for the flow control application of air-supplement PSJs.

Investigation of the aerodynamic characteristics of horizontal axis wind turbine using an active flow control method via boundary layer suction

The boundary layer suction (BLS) has the merit of effectively suppressing the formation and development of separation vortices by removing low-momentum fluid occurs on the leading-edge of airfoil suction surface. The present paper numerically investigates the impacts of BLS on the aerodynamic performance of a horizontal axis wind turbine (HAWT). First, the dynamic mode decomposition method is adopted to gain an insight into the improvement mechanism of BLS control by analyzing the flow patterns of a two-dimensional S809 airfoil. Subsequently, the BLS is arranged on a NREL Phase VI reference turbine to study the power increment and flow field characteristics, and the effects of different working angles αs and momentum coefficients Cμ are investigated at different tip speed ratios (TSRs). The results indicate that the use of BLS control can increase the pressure difference between the suction surface and the pressure surface, raise the peak negative pressure at the leading edge and shift the airfoil flow separation point backwards, and thus effectively enhance the stability of the flow field. Furthermore, the BLS control can significantly improve the power extraction at low to medium TSR if this active device is appropriately implemented on the HAWT.

Performance improvement of adaptive flap on flow separation control and its effect on VAWT

Flow separation on blades is the main factor to reduce the power efficiency of vertical axis wind turbine (VAWT). A bio-inspired flow control device called adaptive flap is considered to have a good adaptability in mitigating the flow separation. The present study investigated the adaptive flap in terms of improving its performance and exploring its effect on VAWT by the Computational Fluid Dynamics (CFD) method combining the SST k-ω turbulence model. Firstly, a static flap with adjustable flap angles was installed on a static airfoil to explore the influence of the flap angle on the flap performance. Then, to improve the performance of the free rotating flap, a linear torque related to the aerodynamic moment acting on the static flap was proposed to constrain the flap. Besides, another torque composited from the aerodynamic moment distribution was proposed to give full control of the flap. The verification results showed that the two torques both worked in two typical airfoil flow separation processes, and the composite torque performed better. Finally, the rotating flaps improved by the composite torque were installed on blades of a VAWT. The simulation results showed that the flap sensibly improved the power coefficient of the VAWT at tip speed ratios (TSRs) of 0.6 and 1.2, providing a solution to slow down the flow separation in WAWT.

Dynamic stall control over an airfoil by NS-DBD actuation**Project supported by the National Natural Science Foundation of China (Grant No. 11802341) and the Open Fund from State Key Laboratory of Aerodynamics of China (Grant No. SKLA20180207).

The wind tunnel test was conducted with an NACA 0012 airfoil to explore the flow control effects on airfoil dynamic stall by NS-DBD plasma actuation. Firstly, light and deep dynamic stall states were set, based on the static stall characteristics of airfoil at a Reynolds number of 5.8 × 105. Then, the flow control effect of NS-DBD on dynamic stall was studied and the influence law of three typical reduced frequencies (k = 0.05, k = 0.05, and k = 0.15) was examined at various dimensionless actuation frequencies (F+ = 1, F+ = 2, and F+ = 3). For both light and deep dynamic stall states, NS-DBD had almost no effect on upstroke. However, the lift coefficients on downstroke were increased significantly and the flow control effect at F+ = 1 is the best. The flow control effect of the light stall state is more obvious than that of deep stall state under the same actuation conditions. For the same stall state, with the reduced frequency increasing, the control effect became worse. Based on the in being principles of flow separation control by NS-DBD, the mechanism of dynamic stall control was discussed and the influence of reduced frequency on the dynamic flow control was analyzed. Different from the static airfoil flow separation control, the separated angle of leading-edge shear layer for the airfoil in dynamic stall state is larger and flow control with dynamic oscillation is more difficult. The separated angle is closely related to the effective angle of attack, so the effect of dynamic stall control is greatly dependent on the history of angles of attack.

Unified mechanisms for separation control around airfoil using plasma actuator with burst actuation over Reynolds number range of 103–106

We conduct large-eddy simulations of separated airfoil flows with control by a dielectric-barrier-discharge plasma actuator over a wide range of Reynolds numbers. The Reynolds numbers based on the chord length (Re) are set at Re = 5.0 × 103, 1.0 × 104, 6.3 × 104, 2.6 × 105, and 1.6 × 106. These Reynolds numbers cover most of the conditions used in the previous studies on separation control by a plasma actuator. The burst frequency nondimensionalized by the chord length and freestream velocity (F+) is used as the computational parameter, and the effective burst actuation and control mechanisms at each Reynolds number condition are investigated. With regard to cases without the control, the flows separate near the leading edge in the laminar state at the Reynolds number range of 103–105, and a substantial turbulent separation occurs at the Reynolds number of 1.6 × 106. Separation control with a high burst frequency [F+ ≃ O(10)] can cause early flow reattachment through the promotion of turbulent transition of a separation shear-layer for Re = 6.3 × 104 and 2.6 × 105. Flow reattachment is mainly caused by momentum entrainment into the boundary layer by fine-scale turbulent vortices. On the other hand, the large-scale spanwise vortices play an important role at F+ = 1 for Re = 1.0 × 104 and 1.6 × 106. In these cases, the dynamics of the spanwise vortices show similar behavior and the pairing of these vortices significantly contributes to the separation control by increasing the momentum entrainment. The optimum value of F+ changes with a Reynolds number. In contrast, when a nondimensional burst frequency based on the characteristics of the separation shear-layer (Fθs) is considered, a high lift-to-drag ratio is found at Fθs≃O(10−2) for all Reynolds numbers. This demonstrates that one of the effective burst frequencies is closely related to the scale of the separation shear-layer, especially for the spanwise vortex shed from the separation shear-layer.

Simulations of Airfoil Suction Side Flow Separation Control Using Plasma Aerodynamic Actuation

Plasma flow control, as a novel active flow control technique, has become a newly-rising focus of international aerodynamics and aerothermodynamics fields. Surface DBD plasma aerodynamic actuation is widely studied in subsonic flow separation control. Numerical simulation investigation of flow stall separation suppression of NACA 0015 airfoil by plasma aerodynamic actuation was performed. The plasma aerodynamic actuation was simplified as momentum and thermal source terms. The numerical model of airfoil flow separation control by plasma aerodynamic actuation was constructed. The simulation results showed that when the inflow velocity was 100m/s and angle of attack was 20 degree, the flow separates at the leading edge of the suction side of the airfoil. There is a 50.3% lift augmentation and 42.0% drag reduction after the plasma aerodynamic actuation was on. The best position of plasma aerodynamic actuation is just near the separation point. The simulation results were basically consistent with the test results, which demonstrated that the simulation model was correct. The results are useful to reveal the mechanism of the correlation of plasma aerodynamic actuation and boundary layer.

Optimal performance of adaptive flap on flow separation control

The stall caused by flow separation of airfoil seriously compromises the efficiency and safety of wind turbines. Inspired by bird feathers, a passive flow control strategy was investigated by configuring adaptive flaps on the suction surface of the airfoil to delay flow separation. The Computational Fluid Dynamics (CFD) software ANSYS Fluent 17.0 using SST k–ω turbulence model was performed to study the performance of the adaptive flap with different length and location. The performance was evaluated by the comparison of total lift coefficient between the flap airfoil and clean airfoil. In addition, the performance of the double flaps configuration was investigated by Fluid-Solid Interaction (FSI) method. The results show that when the separation point is at the upstream of flap hinge point, the total lift coefficient increases with the flap rising, and when the flap reaches the optimal flap angle, the lift coefficient begins to decrease. When the flap operates in optimal angle, the closer the flap is to the trailing edge of the airfoil, the more the lift enhancement for moderate flow separation, and the closer the flap is to the leading edge of the airfoil, the more the lift enhancement for serious flow separation. For the flaps with the same location, the longer flap can block more backflow resulting in a better performance, but when the flap is long enough to block the fluid, the shorter flap has better performance. Considering the variations of the flap aerodynamic moment with the flap angle, an external hinge moment was recommended to improve the swinging flap performance.

Airfoil flow separation control with plasma synthetic jets at moderate Reynolds number

An array of 26 plasma synthetic jet actuators (PSJA) is flush-mounted on a NACA-0015 airfoil model to control the leading-edge flow separation at moderate Reynolds number (Re_c=1.7times 10^5). The stall angle of this airfoil is postponed from 15.5^circ to approximately 22^circ, and the peak lift coefficient is increased by 21%. PSJAs exhibit distinctive separation control mechanisms depending on the relative location between actuation and separation and reduced frequency of actuation (F^*). At an angle of attack of alpha =15.5^circ, the non-actuated flow separates approximately 4% chord length downstream of the jet orifices . Plasma synthetic jets (PSJs) applied at F^*ge 0.5 can displace the separation point downstream to mid-chord position, as a result of the energizing of the incoming boundary layer through mixing enhancement. As a comparison, with actuation frequency of F^*le 0.25, the separation point at alpha =15.5^circ remains near the leading edge and the zero-velocity line is periodically swept towards the suction surface by the convecting spanwise vortices generated from PSJ actuation, leading to a reduction of time-averaged backflow area. For the case of separation control at alpha =22^circ, the separation point resides always upstream of the actuation position, regardless of actuation frequency. The peak lift coefficient is attained at F^*=1, and the decreasing lift at high actuation frequency (F^*=2) is ascribed to the severe interaction between adjacent spanwise vortices at short spacing.Graphical abstract

A combination application of tandem blade and endwall boundary layer suction in a highly loaded aspirated compressor outlet vane

Aspirated compressor is a promising design concept to enhance the power density of the compression system; however, with regard to the rear stages of multistage aspirated compressor, the blade is fairly thin. Limited by the mechanical constraints, it seems impractical to implement the boundary layer suction on the blade suction surface. So the question arising is can we replace the blade suction surface with other feasible flow control methods without implementing extra device on the blade? To address this issue, a compound flow control method, composed of the endwall boundary layer suction and tandem blade, is proposed. The design philosophy is to utilize the EBLS to suppress the three-dimensional corner stall while to use the tandem blade to control the two-dimensional airfoil flow separation. The endwall boundary layer suction is barely implemented in the forward blade, whereas the corner flow in the rear blade is restrained by the flow through the gap between the forward and rear blades. The preliminary implement strategy of the compound flow control was presented and then applied in the design of a highly loaded aspirated compressor outlet vane. Three-dimensional numerical simulations were carried out to validate its effectiveness with different inlet boundary layer distributions. Both flow fields in the outlet vane and its loss characteristics were analyzed. The results show that, by applying the compound flow control, the outlet vane could not only achieve an aggressive loading without incurring large-scale separation at the design point but also have a considerable available incidence range. Due to the implement of the endwall boundary layer suction, the tandem blade can bring out its full potential in the two-dimensional flow control. Moreover, owing to the flow through the gap of the forward and rear blades, the aspiration flow rate required for the suppression of the three-dimensional corner stall can be reduced.

Main characteristics of suction control of flow separation of an airfoil at low Reynolds numbers

The suction control of flow separation of a NACA0012 airfoil is studied numerically for Reynolds number Re=104, angle of attack α=2°, 4°, 6°, and 8°, and Re=105, α=4°. The suppression effects of suction are studied considering many parameters such as suction coefficient, location, angle, hole width, and porosity. The variations in energy consumption and lift–drag ratio during the control process are used for evaluating the control effects.

Effect of 3D stall-cells on the pressure distribution of a laminar NACA64-418 wing

A 3D stall-cell flow-field has been studied in a 4.8 aspect-ratio wing obtained by linear extrusion of a laminar NACA64-418 airfoil profile. The span-wise change in the velocity and pressure distribution along the wing has been quantified with respect to the development of cellular structures from $$8^{\circ }$$ to $$20^{\circ }$$ angle of attack. Oil-flow visualizations help localizing the regular cellular pattern in function of the angle of attack. Multi-plane stereoscopic PIV measurements obtained by traversing the entire setup along the wing span show that the flow separation is not span-wise uniform. The combination of different stereoscopic fields into a 3D volume of velocity data allows studying the global effect of the stall-cell pattern on the wing flow. Integration of the experimentally computed pressure gradient from the Navier–Stokes equation is employed to compute the span-wise distribution of the mean surface pressure. Comparison of the results with the ones obtained from pressure taps installed in the wing evidences a span-wise periodic loading on the wing. The periodic loading has maxima confined in the stream-wise direction between the location of the highest airfoil curvature and the one of the airfoil flow separation. Estimation of the periodic loading is found within 2–6 % of the sectional wing lift.

Geometrical effects on the airfoil flow separation and transition

We present results from direct numerical simulations (DNS) of incompressible flow over two airfoils, NACA-4412 and NACA-0012-64, to investigate the effects of the airfoil geometry on the flow separation and transition patterns at Re=104 and 10 degrees incidence. The two chosen airfoils are geometrically similar except for maximum camber (respectively 4%C and 0 with C the chord length), which results in a larger projection area with respect to the incoming flow for the NACA-4412 airfoil, and a larger leeward surface curvature at the leading edge for the NACA-0012-64 airfoil. The governing equations are discretized using an energy conservative fourth-order spatial discretization scheme. An assessment on the two-point correlation indicates that a spanwise domain size of 0.8C is sufficiently large for the present simulations. We discuss flow separation at the airfoil leading edge, transition of the separated shear layer to three-dimensional flow and subsequently to turbulence. Numerical results reveal a stronger adverse pressure gradient field in the leading edge region of the NACA-0012-64 airfoil due to the rapidly varying surface curvature. As a result, the flow experiences detachment at x/C=0.08, and the separated shear layer transition via Kelvin–Helmholtz mechanism occurs at x/C=0.29 with fully developed turbulent flow around x/C=0.80. These flow development phases are delayed to occur at much downstream positions, respectively, observed around x/C=0.25, 0.71 and 1.15 for the NACA-4412 airfoil. The turbulent intensity, measured by the turbulent fluctuations and turbulent Reynolds stresses, are much larger for NACA-0012-64 from the transition onset until the airfoil trailing edge, while turbulence develops significantly downstream of the trailing edge for the NACA-4412 airfoil. For both airfoils, our DNS results indicate that the mean Reynolds stress u′u′‾/U02 reaches its maximum value at a distance from the surface approximately equal to the displacement thickness, consistent with the experimental observations (Boutilier & Yarusevych, Phys. Fluids, 2012). A quantitative eigen-system analysis on the instantaneous velocity field shows that although the flow over an airfoil is intrinsically anisotropic, the alignments between the vorticity vector and the eigenvectors ofSij and SikSkj+ΩikΩkj are quite similar to those of the homogeneous isotropic turbulent flows due to the formation of vortex tubes.

Unsteady Flow Separation and High Performance of Airfoil with Local Flexible Structure at Low Reynolds Number

AbstractThe unsteady flow separation of airfoil with a local flexible structure (LFS) is studied numerically in Lagrangian frames in detail, in order to investigate the nature of its high aerodynamic performance. For such aeroelastic system, the characteristic-based split (CBS) scheme combined with arbitrary Lagrangian-Eulerian (ALE) framework is developed firstly for the numerical analysis of unsteady flow, and Galerkin method is used to approach the flexible structure. The local flexible skin of airfoil, which can lead to self-induced oscillations, is considered as unsteady perturbation to the flow. Then, the ensuing high aerodynamic performances and complex unsteady flow separation at low Reynolds number are studied by Lagrangian coherent structures (LCSs). The results show that the LFS has a significant influence on the unsteady flow separation, which is the key point for the lift enhancement. Specifically, the oscillations of the LFS can induce the generations of moving separation and vortex, which can enhance the kinetic energy transport from main flow to the boundary layer. The results could give a deep understand of the dynamics in unsteady flow separation and flow control for the flow over airfoil.

Flow separation control over a Gö 387 airfoil by nanosecond pulse-periodic discharge

Airfoil flow separation control using plasma actuator driven by repetitive nanosecond pulse voltage was experimentally investigated. The pressure distribution on an airfoil surface was measured by means of a liquid manometer, which is free from electromagnetic interference of the plasma. By integrating the pressure distribution, the lift coefficient was computed and the effects of the input voltage amplitude and repetitive frequency were evaluated. The results show two different manners of the lift increment depending on the angle of attack. At the pre-stall and stall angle, the flow is steady and the lift increment does not depend on the frequency. A strong hysteresis effect is also observed, i.e., once the lift increases due to the plasma actuation, it is still increased even after the actuation stops. At the post-stall angle, the flow is unsteady and the lift increment becomes significant with actuation at frequencies related to inherent flow instabilities for the shear layer and wake, which are determined by the spectrum analysis based on hot-wire measurements. At the preference frequency ranges, there is a certain voltage amplitude depending on the angle of attack, at which the extended plasma layer results in an optimum increase of lift.

Numerical Analysis of Role of Bumpy Surface to Control the Flow Separation of an Airfoil

The efficiency of mission varying aircrafts can be altered by either flow control method or adaptive wing technology. Here we have researched the control of the flow separation of airfoil by providing partial bumpy on the upper surface of the geometry. Flow of low speed is considered. For the analysis purpose NACA 4315 was chosen. Two models were generated (i) regular airfoil using NACA 4315 profile (ii) providing bumpy surface on the NACA 4315 surface on the trailing edge at 80%C. Numerical approach is undertaken to observe the flow separation on the aerofoil. From the observation it was noted that by using the bumpy surface on the upper surface of the aerofoil the flow separation was delayed. Flow separation occurs at 9 degree angle of attack in the smooth surface whereas in bumpy surface it occurs at 15 degree angle of attack. Thus, indicating the increase of lift force and control of flow separation of an airfoil due to bumpy surface.

Numerical Investigation of the Controllable Wing Stall Caused by the Air Injection

This paper presents results of numerical investigation on a controllable airfoil flow separation phenomena practically applied in Formula One racing by a device called the F-duct. Separation is forced by air jets from slots located at different positions on the profile of the dual element wing and is intended to reduce aerodynamic drag. Slot position and the air jet velocity are the main parameters controlling the flow separation. The flow structure, surface pressure distribution, and the generated downwards lift and drag forces were investigated in this study. Two different flow separation structures have been recognised. Typically, wing stall is correlated with an increase in aerodynamic drag force. However, in the case of the finite wing with low aspect ratio, the induced drag is dominant and is proportional to the downforce. Therefore, flow separation on the wing increases the profile drag while simultaneously reducing the induced drag, resulting in a decrease in the total aerodynamic drag.