Rounded Trailing Edge Research Articles

CFD Investigation of Dual Synthetic Jets on an Optimized Aerofoil's Trailing Edge

In fluid dynamics, a flow control device is used to control, manage, or modify the behavior of a fluid flow. Jet actuators work by releasing high-velocity jets of fluid, usually air or gas, into the surrounding environment to control or manipulate the flow of fluids. In this study, the flow control device, which was a dual synthetic jet actuator (DSJA), acted as a lift enhancement device over an optimized NACA 0012 aerofoil with a rounded trailing edge (TE) (Coanda surface approximately 9% of the trailing edge was modified) to enhance the lift at various angles of attack (AOAs). Fluctuating pressure inlets were introduced in two slots. When the dual synthetic jets were in control, the out-of-phase jets from the upper and lower trailing edge jets helped to boost the lift coefficient. The suction stroke from the lower half of the jet made the Coanda effect stronger in the upper half. The upper trailing edge jet deflected downwards merged with the lower one and helped to deflect the flow field closer to the bottom half. An unsteady CFD analysis was performed on optimized airfoils with and without a DSJ, with a driving frequency of 40.6 and a reduced frequency of 0.025 at a Reynolds number of 25000. The results obtained at different angles indicated that the L/D ratio was improved by 13.5% at higher angles of attack in the presence of the DSJA.

Numerical investigation of elliptical trailing edge on cascade loss of transonic turbine airfoils

Abstract The shape and thickness of the trailing edge are important factors in the supersonic loss in transonic cascades. To investigate the influence of trailing edge shape, a prismatic turbine cascade and a three-dimensional cascade with an elliptic trailing edge were analyzed with numerical simulation by changing different elliptic axis ratios ER. It was indicated that the shock wave and its mixing with the wake were the important sources of loss in the transonic prismatic cascade. With the increase of exit Mach number, shock waves are generated rapidly and migrate downstream, and the shock loss and the total loss of the prismatic cascade are both accrescent. Increasing the elliptical axis ratio could weaken shock waves and delay shock waves. In the loss state region 3, the cascade loss of the ER3 blade is reduced by 8% compared with the round trailing edge cascade. As the elliptical axis ratio of the three-dimensional cascade is increasing, the profile loss and secondary loss are reduced. However, an excessive elliptical axis ratio could lead to an increase in tip leakage loss. Within the range of ER values from 0 to 3, the total loss of the transonic three-dimensional cascade reached minimization when ER was 3.0.

Fluidic Flow Control Devices for Gust Load Alleviation

A reduction in the structural weight and climate-relevant emissions of future transport aircraft can be realized through active gust load alleviation, limiting the peak aerodynamic loads experienced in flight. Two fluidic concepts, surface and Coandă jets, offer promising solutions. The surface jet induces a separated flow region on the wing’s suction side near the trailing edge, while the Coandă jet utilizes tangential blowing over a rounded trailing edge to make use of the Coandă effect for direct circulation control. This study investigates these fluidic actuation concepts using two-dimensional unsteady Reynolds-averaged Navier–Stokes simulations on a supercritical airfoil to alleviate gust-induced lift increases. Results reveal that fluidic concepts can surpass the load reduction capability of a trailing edge flap. Notably, the Coandă-type design, while more efficient, exhibits limited control authority compared to surface jet, achieving maximum gust load reductions of 40% versus 70% for surface jet. Optimizing the temporal deployment of the surface jet leads to increased peak lift reduction. Even under challenging conditions with reduced gust anticipation time and actuator placement near the wing tip with a reduced chord length, the surface jet maintains consistent performance and therefore offers a promising solution for active gust load alleviation on future aircraft.

Numerical Investigations on Unsteady Features of Rounded Trailing Edge Airfoils

As a practical rounded trailing edge airfoil for coaxial rotors, DBLN-526 is a fore and aft symmetrical airfoil with two steps on its lower side. This airfoil has been used at the inboard section of the coaxial rotor system. As there are always two eddies behind the airfoil because of its rounded trailing edge, the interaction between the separation and transition should be considered. Thus, the unsteady Reynolds-averaged Navier-Stokes- (RANS-) based γ − Re ¯ θ t model was used to analyze the unsteady transition and separation features. Three different rounded trailing-edge airfoils were compared with DBLN-526. Power spectrum density analysis and Δ Cl calculations demonstrated that the lift coefficient fluctuation of DBLN-526 was smaller than that of other rounded airfoils with different angles. Further investigation indicated that the locations of transition for DBLN-526 can be fixed at a wide range of angles by the unique design on its lower side. Because of this settled transition location, the size of separation is decreased, and the position of separation is settled as well, which leads to a lower lift coefficient fluctuation. The turbulent kinetic energy after the transition was higher, which injected a lot of energy into the boundary layer, and the separation zone near the transition position was relatively smaller. This study provides an indication for controlling separation and reducing unsteady fluctuations for rounded trailing edge airfoils.

Lift enhancement based on virtual aerodynamic shape using a dual synthetic jet actuator

The Dual Synthetic Jet Actuator (DSJA) is used to develop a new type of lift enhancement device based on circulation control, and to control the flow over the two-dimensional (2D) NACA0015 airfoil. The lift enhancement device is composed of a DSJA and a rounded trailing edge (Coanda surface). The two outlets of the DSJA eject two jets (Jet 1 and Jet 2). Jet 1 ejects from the upper trailing edge, which increases the circulation of airfoil with the help of the Coanda surface. Jet 2 ejects from the lower trailing edge, which acts as a virtual flap. The Reynolds number based on the airfoil chord length and free flow velocity is 250000. The results indicate that the circulation control method based on Dual Synthetic Jet (DSJ) has good performance in lift enhancement, whose control effect is closely related to momentum coefficient and reduced frequency. With the increase of the reduced frequency, the control effect of the lift enhancement is slightly reduced. As the momentum coefficient increases, the control effect becomes better. When the angle of attack is greater than 4°, the increments of lift coefficients under the control of DSJ are larger than those under the control of the steady blowing at a same momentum coefficient. The maximum lift augmentation efficiency can reach 47 when the momentum coefficient is 0.02, which is higher than the value in the case with steady blowing jet circulation control.

Experimental Study of a NACA0015 Circulation Control Airfoil Using Synthetic Jet Actuation

Wind tunnel testing is conducted on a NACA0015 airfoil with a rounded trailing edge, known as a circulation control airfoil, at the chord Reynolds number of and at the angles of attack of 0–15°. An array of synthetic jet actuators is installed next to the rounded trailing edge to produce synthetic jet at the excitation frequency range of 150–600 Hz and the momentum coefficient range of 0.00021–0.0134 across the airfoil span length. For the circulation control airfoil at prestall angles of attack, the current study suggested that the reduced excitation frequency of a synthetic jet should approximately be on an order of the Strouhal number based on the diameter of the Coanda surface to achieve optimal lift augmentation.

Spatiotemporal correlation and control of wall pressure fluctuations induced by subsonic cavity flows

Strong wall pressure fluctuations are generated mostly by a high-speed flow passing over a cavity and should be accurately predicted and controlled for the structural design of aircraft. In this work, a spatiotemporal analysis was performed to estimate wall pressure fluctuations and reveal the mechanism of flow control by a sawtooth-like vortex generator installed on the leading edge or by the use of a rounded trailing edge in suppressing wall pressure fluctuations. Both the cavity tones and overall sound pressure level were analyzed under different flow speeds and flow control approaches. To investigate the power transfer in the temporal and spatial domains, power spectral density contours were drawn and discussed. For the mode separation, the phased array technique with an improved beamforming algorithm was adopted to calculate the wavenumber map and then to separate the acoustic modes from the convective modes by utilizing the spatial correlation of the wall pressure fluctuations. The mode separation results show that the convective modes contribute mainly to the low-frequency pressure fluctuations, while the acoustic modes contribute to both the low-frequency and high-frequency pressure fluctuations. Flow control by the rounded trailing edge approach is more effective than that by the serrated structure approach in reducing the cavity acoustic environment, which enables wall pressure fluctuations to transfer from the high-frequency range to the low-frequency range compared with the “no control” case.

Hydrodynamic analysis of an underwater glider wing using ANSYS fluent as an investigation tool

The analysis of the hydrodynamic behavior of the modified trailing edge of an underwater glider wing was conducted using ANSYS Fluent 18.1. The procedure was carried out using the computational fluid dynamics (CFD) approach. A standard four digits NACA0016 aerofoil profile was selected while the geometry points were extracted from NACA online aerofoil generator using Solidworks, which enabled the creation of the 3D wing profile. Since the focus of the study is more toward hydrodynamic behavior or the influence of the modified round trailing edge as opposed to the normal sharp trailing edge, coefficient of lift (CL), coefficient of drag (Cd), coefficient of pressure (Cp) and the drag polar (Cd/CL) are very crucial parameters and have an influence on the accuracy of the results. The results showed lift and drag coefficient are a function of the angle of attack (AOA) and free stream velocities. The investigation showed the optimal design point of the AOA of 12° at the corresponding speed of 0.26 m/s. The critical AOA matched with the optimal design point of 12°. The results also showed that Cp is higher closer to the fuselage and therefore explains why more structural rigidity must be provided in this region of high stress concentration. The drag and lift curve corresponded to the typical wing characteristics of the published experiment data regardless of the operating conditions.

Low Reynolds airfoil family for small horizontal axis wind turbines based on RG15 airfoil

The present study introduces a low Reynolds number (Re) airfoil family for the entire blade span of small wind turbines, aiming to reduce the effects related to laminar separation, improve startup response and meet acceptable levels of structural integrity. Six airfoils of varying relative thickness were designed by increasing the thickness distribution of RG15 airfoil up to 50% and adopting a rounded trailing edge with a diameter equal to 1% of the chord length. The aerodynamic performance of RG15 family was initially evaluated by means of XFOIL code at several low Reynolds numbers ranging from 60,000 to 300,000 and angles of attack between − 6° and 14°. XFOIL analysis revealed that increasing relative thickness leads to the reduction of maximum lift-to-drag ratio $$(C_{L} /C_{D} )$$ . However, this percentage reduction decreases with increasing Re. The maximum reduction in maximum $$C_{L} /C_{D}$$ was found at 60,000 Re for the thickest airfoil of RG15 family. Conversely, a growth of maximum lift coefficient $$(C_{L} )$$ was observed by increasing relative thickness. Besides, a Reynolds-Averaged Navier–Stokes (RANS) analysis was also conducted at 300,000 Re to get additional information on the flow characteristics. Comparisons between the XFOIL and RANS data were performed. The results of RANS simulations were generally in accordance with those of XFOIL. However, notable over-estimations of drag coefficient $$(C_{D} )$$ were detected. The behavior of the recirculation area behind the rounded trailing edge and that of the separation bubble near the leading edge for different values of relative thickness and angle of attack was examined. Thickening of the airfoils was found to have a beneficial impact on the appearance of separation bubbles, while no significant effect of the rounded trailing edge on $$C_{L}$$ and $$C_{D}$$ was observed.

Numerical simulation of flow past a flat plate with anti-singing trailing edge

Aerofoil with anti-singing trailing edge has been seen as the treatment to eliminate singing. The present study numerically investigates the bevel and chamfer anti-singing trailing edges with rounded ones in flat plates. The effects of 25°, 45°, 10° bevel and 15°, 45° chamfer angles provided at the trailing edge of the flat plate on the aerodynamic characteristics are systematically compared with rounded ones at a Reynolds number of 1.5x105 and 25°, 30° and 45° angle of attack. The substantial reduction of primary vortex shedding frequency is observed in 25°, 45°, 10° bevel and 15°, 45° chamfer trailing edges as compared to the rounded one. Also, the presence of broader wake profile indicates higher drag in rounded trailing edge geometry as compared to anti-singing trailing edges. Thus, the present study demonstrates that the 45° and 10° bevel trailing edge could provide better anti-singing behavior and aerodynamic performance in comparison with 25° bevel, 15°, 45° chamfer and rounded trailing edge.

Wave drag on asymmetric bodies

An asymmetric body with a sharp leading edge and a rounded trailing edge produces a smaller wave disturbance moving forwards than backwards, and this is reflected in the wave drag coefficient. This experimental fact is not captured by Michell’s theory for wave drag (Michell Lond. Edinb. Dubl. Phil. Mag. J. Sci., vol. 45 (272), 1898, pp. 106–123). In this study, we use a tow-tank experiment to investigate the effects of asymmetry on wave drag, and show that these effects can be replicated by modifying Michell’s theory to include the growth of a symmetry-breaking boundary layer. We show that asymmetry can have either a positive or a negative effect on drag, depending on the depth of motion and the Froude number.

Understanding the secondary separation from an inclined square cylinder with sharp and rounded trailing edges

The steady separated flow past a square cylinder at 45° incidence and having a sharp/rounded base is a unique kind of separated flow around a symmetric bluff obstacle. At low Reynolds numbers (Re ≈ 6–8), this flow displays diverse features that are not known to be displayed by common symmetric bluff bodies, such as circular/square cylinders at zero incidence. The uniqueness of this flow was first highlighted by Kumar et al. [“Steady separation of flow from an inclined square cylinder with sharp and rounded base,” Comput. Fluids 171, 29–40 (2018)] in their recent numerical investigation concerning the onset of initial separation. They reported detailed flow kinematics resulting from separation of the laminar boundary layer via secondary mode, primary mode, and their combination. In this work, based on stabilized finite-element computations, the fluid dynamic aspects of primary and secondary separation are delved. The parameters considered are the normalized corner radius at the cylinder base and Re. Exploration of the influence of base radius on the occurrence/absence of the secondary separation is the main objective. The occurrence of secondary separation is a function of base radius. It occurs only with small or no rounding at the base, i.e., when the pressure gradient close to the base is favorable or mildly adverse. Irrespective of the base radius, the onset of secondary separation occurs at a constant Re value of 7.3. The drag decays as Re−0.62; it becomes invariant when Re is held constant. A new nondimensional parameter η accommodating the effects of pressure recovery and radius of curvature is introduced. The drag exhibits no sensitivity to η.

The influence of reduced Reynolds number on the wake of the DrivAer estate vehicle

Experiments on vehicle aerodynamics were conducted in a low-speed water channel at a Reynolds number two orders of magnitude lower than that of typical full-scale studies. Particle Image Velocimetry (PIV) measurements, focused on the wake of a realistic estate passenger vehicle model, were obtained and compared with existing wake measurements from full-scale wind-tunnel tests. The combination of sharp and rounded trailing edges present on the geometry impacts properties of the separating shear layers, although the general wake structures are qualitatively similar. In both cases, the wake remains dominated by the upwash from the underbody exit diffuser outflow. Effects of the reduced Reynolds number on the spectral content of the wake velocity field downstream of the recirculation region are also documented.

The Role of Vortex Shedding in the Trailing Edge Loss of Transonic Turbine Blades

The loss of square, round, and elliptical turbine trailing edge geometries, and the mechanisms responsible, is assessed using a two-part experimental program. In the first part, a single blade experiment, in a channel with contoured walls, allowed rapid testing of a range of trailing edge sizes and shapes. In the second part, turbine blade cascades with a subset of sizes of the trailing edge geometries tested in part one were evaluated in a closed-loop variable density facility, at exit Mach numbers from 0.40 to 0.97, and exit Reynolds numbers from 1.5 × 105 to 2.5 × 106. Throughout the test campaign, detailed instantaneous Schlieren images of the trailing edge flows have been obtained to identify the underlying unsteady mechanisms in the base region. The experiments reveal the importance of suppressing transonic vortex shedding, and quantify the influence of this mechanism on loss. The state and thickness of the blade boundary layers immediately upstream of the trailing edge are of critical importance in determining the onset of transonic vortex shedding. Elliptical trailing edge geometries have also been found to be effective at suppressing transonic vortex shedding. For trailing edges that exhibit transonic vortex shedding, a mechanism is identified whereby reflected shed shockwaves encourage or discourage vortex shedding depending on the phase with which the shocks return to the trailing edge, capable of modifying the loss generated.

Experimental and Numerical Study of Transonic Cooled Turbine Blades

State-of-the-art gas turbines (GT) operate at high temperatures that exceed the endurance limit of the material, and therefore the turbine components are cooled by the air taken from the compressor. The cooling provides a positive impact on the lifetime of GT but has a negative impact on its performance. In convection-cooled turbine blades the coolant is usually discharged through the trailing edge and leads to limitations on the minimal size of the trailing edge, thereby negatively affecting the losses. Moreover, the injection of cooling air in the turbine disturbs the main flow, and may lead to an additional increase in loss. Trailing edge loss is a significant part of the overall loss in modern gas turbines. This study comprises investigations of the unguided flow angle, the trailing edge shape, and cooling air injection through the trailing edge on the base pressure and profile losses in cooled blades. Some vane and blade cascades with different unguided turning angle and two shapes of trailing edges with and without coolant injection were studied both experimentally and numerically. This analysis provides a split of losses caused by different factors, and offers opportunities for efficiency and lifetime improvements of real engine designs/upgrades. In particular, it is shown that an increase in the unguided turning angle and the use of a round trailing edge result in a reduction of loss in case of a relatively thick trailing edge. Numerical investigation showed that an increase in the unguided turning angle at the initial transonic vane with a thick and blunt trailing edge, without a change in other basic geometric parameters, allowed for a significant reduction of the profile loss by about 3–4% at the exit Mach number M2is = 0.7–1.0. Experimental investigation of four cascades with cooling air injection into the base flow through the trailing edge allowed us to validate the fact that in blades with a low level of base pressure Cpb < −0.1 at m¯ = 0 a non-monotonic dependence of the change of losses against relative cooling air mass flow m¯ is observed. Firstly, the cooling air injection into wake increases base pressure and decreases losses; then the losses start to increase with increasing cooling mass flow due to the interaction between the main flow and the cooling air (mixing losses) and, finally, due to the cooling mass flow increase and momentum increase losses are decreased. In blades with an increased level of the base pressure coefficient Cpb ≥ −0.1 at m¯ = 0 the cooling air injection results in an increase in losses right from the beginning of the injection and then, according to the cooling mass flow increase and momentum rise, losses decrease. It is also shown that injection through the trailing edge slot parallel to the main flow leads to a neutral loss impact and even a loss reduction in the subsonic range and a loss increase in the supersonic range of exit Mach numbers.

Discontinuity of lift on a hydrofoil in reversed flow for tidal turbine application

This work presents an experimental investigation of a hydrofoil in reversed flow configuration in the context of marine current turbine development. Experiments consist in hydrodynamic force measurements and PIV flow observations on a NACA 0015 hydrofoil, at 5×105 Reynolds number. The hydrofoil in reversed flow produces a higher lift than in the classical forward flow for very low angles of attack and proved to be relatively efficient for an angle of attack lower than 10°, despite a much higher drag than the same foil in direct flow. Moreover, the lift coefficient shows a discontinuity with an hysteresis effect when the angle of attack is varied up and down around zero-degree. It is shown that the sharp leading edge generates an early Leading Edge Separation Bubble on one side (suction side) even for vanishing angles of attack. This separation bubble triggers the transition to turbulence of the boundary layer on the suction side while the pressure side boundary layer remains laminar. As a consequence, separation on the rounded trailing edge occurs farther downstream on the (turbulent) suction side compared to the (laminar) pressure side. The Leading Edge Separation Bubble and the inherent up–down asymmetry in the boundary layer regime are responsible for the lift singularity.

Comparison of flow modification induced by plasma and fluidic jet actuators dedicated to circulation control around wind turbine airfoils

In order to reduce the aerodynamic load fluctuations on wind turbine blades by innovative control solutions, strategies of active circulation control acting at the blade airfoil trailing edge are studied, allowing lift increase and decrease. This study presents a comparison of results obtained by performing surface plasma and continuous fluidic jet actuation on a blade airfoil designed with a rounded trailing edge. In the present study, both actuator types are located at the trailing edge. Plasma actuators act uniformly in the spanwise direction, whereas fluidic jets blow through small squared holes distributed along the span, and therefore, provide a three-dimensional action on the flow. Load and velocity field measurements were performed to assess the effectiveness of both actuators and to highlight the flow mechanisms induced by both actuation methods for lift-up configurations. Results are presented for a chord Reynolds number of 2. 105 and for a lift coefficient increase of 0.06.

Suppression of force fluctuations in flow past an aerofoil

Force fluctuations on a solid body are associated with unsteadiness in the wake, e.g. vortex shedding. Therefore, the control of force fluctuations can be realised by suppressing the flow unsteadiness. A NACA0024 aerofoil closed with a round trailing edge is chosen to represent the solid body throughout this investigation, with the Reynolds number fixed at Re = 1000 and angle of attack α ≤ 15o, at which the uncontrolled flow is two-dimensional. A linear optimal control is calculated by analysing the distribution of sensitivity of unsteadiness to control around the entire surface of the body. The nonlinear effects of the calculated control, which can be actuated through surface-normal suction and blowing across the surface of the aerofoil, are tested through two-dimensional direct numerical simulations. It is observed that a surface-normal velocity control with a maximum magnitude less than 8% of the free stream velocity completely suppresses unsteadiness at α = 10° with an overall drag reduction of 14% and a 138% increase of lift.

Influence of circulation on a rounded-trailing-edge airfoil using plasma actuators

An experimental study on influence of circulation around a symmetric airfoil with a rounded trailing edge is presented. Flow control is achieved by the use of dielectric barrier discharge plasma actuators placed at the trailing edge of the airfoil. Direct lift and drag measurements are taken using an external load balance at freestream velocities of 10, 15 and 20 m/s corresponding to chord Reynolds number of 140,000, 210,000, 280,000. Additionally, time-resolved particle image velocimetry is used in order to elucidate the topology and dynamical response of the wake flow under the influence of actuation. Results indicate an increase in lift coefficient of approximately 0.1 for the lowest tested Reynolds number using the plasma actuator. Flowfield measurements indicate the successful manipulation of the Kutta condition enabled by the plasma actuator. The actuator is enhancing the mixing of the wake near the trailing edge while reducing the dominant shedding frequency. Proper orthogonal decomposition analysis reveals further details regarding the dynamics of the wake flow in presence of actuation, suggesting the sensitivity of the control concept to the positioning of the actuator as well as the angle of attack.

English

In this paper we investigated numerically the dynamic stall phenomenon and the possibilities to control it, with application to vertical axis wind turbines (for urban users). The Phenomenon appear at low tip speed ratio (TSR<4) and it has a great impact on structural integrity of the wind turbine and power performances. For this reason we performed a computational study of dynamic stall around NACA 0012 airfoil in pitching motion at relative low Reynolds number (10 5 ). Also, we performed the same analysis for four flow control methods: two passive (Gurney flap and slot) and two active (blowing jet on the rounded trailing edge and synthetic jet periodically activated). The Results are compared to those of an existing experimental case test.