Trailing Edge Flow Research Articles

Trailing-edge flow and noise control using porous treatments

This paper is concerned with the application of porous treatments as a means of flow and aerodynamic noise reduction. An extensive experimental investigation is undertaken to study the effects of flow interaction with porous media, in particular in the context of the manipulation of flow over blunt trailing edges and attenuation of vortex shedding. Comprehensive boundary layer and wake measurements have been carried out for a long flat plate with solid and porous blunt trailing edges. Unsteady velocity and surface pressure measurements have also been performed to gain an in-depth understanding of the changes to the energy–frequency content and coherence of the boundary layer and wake structures as a result of the flow interaction with a porous treatment. Results have shown that permeable treatments can effectively delay the vortex shedding and stabilize the flow over the blunt edge via mechanisms involving flow penetration into the porous medium and discharge into the near-wake region. It has also been shown that the porous treatment can effectively destroy the spanwise coherence of the boundary layer structures and suppress the velocity and pressure coherence, particularly at the vortex shedding frequency. The flow–porous scrubbing and its effects on the near-wall and large coherent structures have also been studied. The emergence of a quasi-periodic recirculating flow field inside highly permeable surface treatments has also been investigated. Finally, the paper has identified several important mechanisms concerning the application of porous treatments for aerodynamic and aeroacoustic purposes, which can help more effective and tailored designs for specific applications.

Boundary layers at the interface of two different shear flows

We present solutions for the boundary layer between two uniform shear flows flowing in the same direction. In the upper layer, the flow has shear strength a, fluid density ρ1, and kinematic viscosity ν1, while the lower layer has shear strength b, fluid density ρ2, and kinematic viscosity ν2. Similarity transformations reduce the boundary-layer equations to a pair of ordinary differential equations governed by three dimensionless parameters: the shear strength ratio γ = b/a, the density ratio ρ = ρ2/ρ1, and the viscosity ratio ν = ν2/ν1. Further analysis shows that an affine transformation reduces this multi-parameter problem to a single ordinary differential equation which may be efficiently integrated as an initial-value problem. Solutions of the original boundary-value problem are shown to agree with the initial-value integrations, but additional dual and quadruple solutions are found using this method. We argue on physical grounds and through bifurcation analysis that these additional solutions are not tenable. The present problem is applicable to the trailing edge flow over a thin airfoil with camber.

Numerical Studies on a NACA0018 Airfoil Blade HAWT with Trailing Edge Jet Flow

This study analyzed an airfoil blade for a horizontal-axis wind turbine (HAWT) with a trailing-edge jet flow design. This design was realized by drilling a hole in the trailing edge of an NACA0018 blade of a conventional HAWT to serve as a pressure injection nozzle. Five inflow wind speeds and three trailing-edge jet flow conditions were examined in the test. The results revealed the efficiency differences between a HAWT with the new jet flow design and conventional HAWTs. The experimental methods employed involved a wind tunnel experiment and a computational fluid dynamics (CFD) simulation. The results revealed that when the inflow wind speed was low, the trailing-edge jet flow accelerated the initiation phase and increased the rotating speed of the HAWT; however, when the inflow wind speed was high, damping occurred and the rotating speed of the turbine blades decreased.

Mitigation of Turbine Vane Shock Waves Through Trailing Edge Cooling

The immense market demand on the high efficiency and lightweight aero-engines results in designs with compact high-pressure turbine stages experiencing supersonic flow field. In supersonic turbines, shocks appear at the vane trailing edges. The interaction of these shock with the neighbouring airfoils and blades on the adjacent rotor row and consequently create considerable amount of losses on the aerodynamic performance of the turbine. Moreover, periodic excitation created by the interaction of the shock waves and the motion of the turbine rotor causes fatigue problems and reduces the lifetime of the engine. Current study aims to alter the vane shock waves through blowing at the trailing edge. In order to characterize the effect of active blowing on the trailing edge flow field, a series of URANS simulations were conducted on OpenFOAM solver platform. Various blowing schemes were simulated over a simplified trailing edge geometry exposed in supersonic flow. The computations were compared in terms of shock intensity, oscillation frequency and exerted pressure forcing over the downstream components. The results showed that unsteady trailing edge blowing were able to modify the fluctuations observed on the shocks by altering the shock intensity, angle and frequency of oscillations. The classification of the wake unsteadiness, i.e. vortex shedding, in terms of trailing edge characteristics were also accomplished through frequency domain analysis of simulations.

Numerical simulation of gurney flaps lift-enhancement on a low reynolds number airfoil

Two-dimensional steady Reynolds-averaged Navier-Stokes (RANS) equations with transition shear stress transport (SST) model were solved to investigate the effects of Gurney flaps on the aerodynamic performance of a low Reynolds number airfoil. This airfoil was designed for flight vehicles operating at 20 km altitude with freestream velocity of 25 m/s. The chord length (C) of this airfoil is 5 m and the corresponding Reynolds number is 7.76×105. Gurney flaps with the heights ranging from 0.25%C to 3%C were investigated. It has been shown that Gurney flaps can enhance not only the prestall lift but also lift-to-drag ratio in a certain range of angles of attack. Specially, at cruise angle of attack (3°), Gurney flap with the height of 0.5%C can increase lift-to-drag ratio and lift coefficient by 1.6% and 12.8%, respectively. Furthermore, the mechanisms of Gurney flaps to improve the aerodynamic performance were illustrated by analyzing the surface pressure distribution, streamlines and trailing-edge flow structure for this low Reynolds number airfoil. Specially, distinguished from some other numerical researches, the flow details such as the laminar separation bubble and transition phenomena for low Reynolds number airfoil with Gurney flaps were investigated and it was found that Gurney flaps can delay the transition onset position at small angles of attack (≤2°). However, with the increase of angles of attack, Gurney flaps will promote the boundary layer transition.

Effects of flow incidence on aerothermal performance of transonic blade tip clearance flows

This paper is a continuation of the previous work and aims to understand how the inlet incidence flows influence the transonic tip leakage loss-generation processes and overtip heat-transfer. Five incidences ranged from −15 to 15°, and two tip gaps of 0.4% and 1.0% of blade span are investigated. Numerical calculations have been carried out by means of the ANSYS-CFX calculation tool, using the k-ε SST turbulence model to study the sensitivities of transonic blade tip aerothermal performance to flow incidences in an RT27a linear turbine-cascade. Flow patterns over the transonic blade tip at design incidence, overtip flow fields and aerodynamics sensitivity to flow incidence, and transonic tip heat transfer characteristics at different incidences are investigated. The computed results show a fair agreement with the rig data and show that the effects of flow incidence on aerodynamics and heat transfer in the front part of blade tips are bigger than that in the rear part of blade tips, and flow incidence effects at 0.4% tip gap is smaller than that at 1.0% tip gap. At various incidences, as the tip gap height increases, the shockwave reflections become more obvious, and it hence reduces the leakage flow. However, the large tip clearance height induces elliptical TLV core structure, and leads to strong mixing with the mainflow. For the blade-tip trailing-edge flow regions, at two blade tip gaps the overtip heat transfer decreases normally as the incidence increases.

The effect of dense gas dynamics on loss in ORC transonic turbines

This paper describes a number of recent investigations into the effect of dense gas dynamics on ORC transonic turbine performance. We describe a combination of experimental, analytical and computational studies which are used to determine how, in-particular, trailing-edge loss changes with choice of working fluid. A Ludwieg tube transient wind-tunnel is used to simulate a supersonic base flow which mimics an ORC turbine vane trailing-edge flow. Experimental measurements of wake profiles and trailing-edge base pressure with different working fluids are used to validate high-order CFD simulations. In order to capture the correct mixing in the base region, Large-Eddy Simulations (LES) are performed and verified against the experimental data by comparing the LES with different spatial and temporal resolutions. RANS and Detached-Eddy Simulation (DES) are also compared with experimental data. The effect of different modelling methods and working fluid on mixed-out loss is then determined. Current results point at LES predicting the closest agreement with experimental results, and dense gas effects are consistently predicted to increase loss.

NACA0012翼型後流の変動流速をトリガーとしたプラズマアクチュエータのFeedback駆動による後縁剥離制御

NACA0012翼型後流の変動流速をトリガーとしたプラズマアクチュエータのFeedback駆動による後縁剥離制御

Trailing-edge flow control for wind turbine performance and load control

This paper reports an investigation into the performance of trailing-edge flow control devices on horizontal axis wind turbines by solving the three dimensional Reynolds averaged Navier-Stokes equations in the rotational framework. The validation case selected for this work is the NREL Phase VI blade with wind tunnel experimental data. The trailing-edge flow control devices studied include microtabs and microjets installed near the trailing-edge of the rotating blade. The divergent trailing-edge is also included in the study as a passive flow control device due to its practical interest. These trailing-edge devices are implemented on the fixed-pitch NREL Phase VI blade, using the original performance and flow characteristics as a benchmark. Both 2D and 3D simulations are carried out in order to investigate the suitability of the 2D blade sectional design analysis and control for the actual 3D rotating blades.

Discrete Vortex Method with Flow Separation Corrections for Flapping-Foil Power Generators

Energy harvesting using flapping foils has attracted significant attention in the last 10 years. In contrast to conventional turbines, which rely on attached flow for high efficiency, a flapping-foil turbine exploits the separation of fluid flow near the nose of the foil to form a leading-edge vortex, resulting in high instantaneous forces. To reduce the computational cost of simulating a flapping foil undergoing high angles of attack (up to 90 deg), a discrete vortex method that incorporates leading-edge flow separation is used in this study. Corrections using semi-empirical functions are introduced to account for the effects of trailing-edge flow separation. Instantaneous lift and power coefficient as well as the mean power coefficient and efficiency from the discrete vortex method with separation corrections are compared against those from a dynamic stall model and computational-fluid-dynamics results obtained by the lattice Boltzmann method. The discrete vortex method provides better agreement with computational-fluid-dynamics simulations compared to the dynamic stall model.

Numerical and experimental investigation of a beveled trailing-edge flow field and noise emission

Efficient tools and methodology for the prediction of trailing-edge noise experience substantial interest within the wind turbine industry. In recent years, the Lattice Boltzmann Method has received increased attention for providing such an efficient alternative for the numerical solution of complex flow problems. Based on the fully explicit, transient, compressible solution of the Lattice Boltzmann Equation in combination with a Ffowcs-Williams and Hawking aeroacoustic analogy, an estimation of the acoustic radiation in the far field is obtained. To validate this methodology for the prediction of trailing-edge noise, the flow around a flat plate with an asymmetric 25° beveled trailing edge and obtuse corner in a low Mach number flow is analyzed. Flow field dynamics are compared to data obtained experimentally from Particle Image Velocimetry and Hot Wire Anemometry, and compare favorably in terms of mean velocity field and turbulent fluctuations. Moreover, the characteristics of the unsteady surface pressure, which are closely related to the acoustic emission, show good agreement between simulation and experiment. Finally, the prediction of the radiated sound is compared to the results obtained from acoustic phased array measurements in combination with a beamforming methodology. Vortex shedding results in a strong narrowband component centered at a constant Strouhal number in the acoustic spectrum. At higher frequency, a good agreement between simulation and experiment for the broadband noise component is obtained and a typical cardioid-like directivity is recovered.

Design and Scaling of Plasma Streamwise Vortex Generators for Flow Separation Control

The design and scaling of plasma streamwise vortex generators in a variable pressure gradient environment are presented. This involves measurements in the boundary layer on the suction surface of a V-22 wing model which provided a generic flowfield that undergoes trailing-edge flow separation as a precursor to lift stall. Experiments were performed at freestream Mach numbers from 0.1 to 0.2, which resulted in chord Reynolds numbers of to . Different streamwise pressure gradients were examined by changing the airfoil angle of attack. Three components of the mean velocity in the boundary layer on the suction side of the wing model were measured using a five-hole pitot probe. These measurements were used to quantify the plasma streamwise vortex generator generated mean streamwise vorticity and the magnitude of the vorticity reorientation term in the vorticity transport equation that was used in the scaling analysis of the plasma streamwise vortex generator performance. The results validate the prediction of the streamwise vorticity production associated with the reorientation of the boundary-layer mean vorticity to scale as , where is the streamwise length of the plasma streamwise vortex generator array, is the external velocity at the crest of the airfoil, is the characteristic mean velocity of the flow over the plasma streamwise vortex generator array, is the change in the streamwise velocity produced by the plasma actuator, and is the spanwise spacing between the plasma streamwise vortex generator electrodes. The criteria for is that the convective timescale be greater than the timescale for vorticity reorientation in the boundary layer, namely, . Both the plasma streamwise vortex generator generated and were found to scale with . An optimally scaled plasma streamwise vortex generator array on the V-22 wing model produced as much as a 40 point drag count reduction at poststall angles of attack.

A Study of Trailing-Edge Losses in Organic Rankine Cycle Turbines

In this paper, vane trailing-edge losses which occur in organic rankine cycle (ORC) turbines are investigated. Experiments are performed to study the influence of dense gas effects on trailing-edge loss in supersonic flows using a novel Ludwieg tube facility for the study of dense-gas flows. The data is also used to validate a computational fluid dynamics (CFD) flow solver. The computational simulations are then used to determine the contributions to loss from shocks and viscous effects which occur at the vane trailing edge. The results show that dense gas effects play a vital role in the structure of the trailing-edge flow, and control the extent of shock and viscous losses.

On the wake flow of asymmetrically beveled trailing edges

Trailing edge and wake flows are of interest for a wide range of applications. Small changes in the design of asymmetrically beveled or semi-rounded trailing edges can result in significant difference in flow features which are relevant for the aerodynamic performance, flow-induced structural vibration and aerodynamically generated sound. The present study describes in detail the flow field characteristics around a family of asymmetrically beveled trailing edges with an enclosed trailing-edge angle of $$25^\circ$$ and variable radius of curvature R. The flow fields over the beveled trailing edges are described using data obtained by particle image velocimetry (PIV) experiments. The flow topology for different trailing edges was found to be strongly dependent on the radius of curvature R, with flow separation occurring further downstream as R increases. This variation in the location of flow separation influences the aerodynamic force coefficients, which were evaluated from the PIV data using a control volume approach. Two-point correlations of the in-plane velocity components are considered to assess the structure in the flow field. The analysis shows large-scale coherent motions in the far wake, which are associated with vortex shedding. The wake thickness parameter $$y_{f}$$ is confirmed as an appropriate length scale to characterize this large-scale roll-up motion in the wake. The development in the very near wake was found to be critically dependent on R. In addition, high-speed PIV measurements provide insight into the spectral characteristics of the turbulent fluctuations. Based on the time-resolved flow field data, the frequency range associated with the shedding of coherent vortex pairs in the wake is identified. By means of time-correlation of the velocity components, turbulent structures are found to convect from the attached or separated shear layers without distinct separation point into the wake.

Vortex structure analysis of unsteady cloud cavitating flows around a hydrofoil

In this paper, time dependent vortex structures are numerically analyzed for both noncavitating and cloud cavitating flows around a Clark-Y hydrofoil with angle of attack [Formula: see text] at a moderate Reynolds number, [Formula: see text]. The numerical simulations are performed using a transport equation-based cavitation model and the large eddy simulation (LES) approach with a classical eddy viscosity subgrid scale (SGS) model. Compared with experimental results, present numerical predictions are capable of capturing the initiation of cavity, growth toward the trailing edge and subsequent shedding process. Results indicate that in noncavitating conditions, the trailing edge vortex and induced positive vortex shed periodically into the wake region to form the vortex street. In cloud cavitating conditions, interrelations between cavity and vortex induce different vortex dynamics at different cavity developing stages. (i) As attached cavity grows, vorticity production is greatly enhanced by the favorable pressure gradient at the leading edge. The trailing edge flow does not have a direct impact on the attached cavity expansion process. Furthermore, the liquid–vapor interface that moves toward the trailing edge enhances the vorticity in the attached cavity closure region. (ii) When the stable attached sheet cavity grows to its maximum length, the accumulation process of vorticity is eventually interrupted by the formation of the re-entrant jet. Re-entrant jet’s moving upstream leads to a higher spreading rate of the attached cavity and the formation of a large coherent structure inside the attached cavity. Moreover, the wavy/bubbly cavity interface enhances the vorticity near the trailing edge. (iii) As the attached sheet cavity breaks up, this large vortex structure converts toward the trailing edge region, which will eventually couple with a trailing edge vortex shedding from the lower surface to form the cloud cavity. The breakup of the stable attached cavity is the main reason for the vorticity enhancement near the suction surface.

Steady/Unsteady Reynolds-Averaged Navier–Stokes and Large Eddy Simulations of a Turbine Blade at High Subsonic Outlet Mach Number

Reynolds-averaged Navier–Stokes (RANS), unsteady RANS (URANS), and large eddy simulation (LES) numerical approaches are clear candidates for the understanding of turbine blade flows. For such blades, the flow unsteady nature appears critical in certain situations and URANS or LES should provide more physical understanding as illustrated here for a laboratory high outlet subsonic Mach blade specifically designed to ease numerical validation. Although RANS offers good estimates of the mean isentropic Mach number and boundary layer thickness, LES and URANS are the only approaches that reproduce the trailing edge flow. URANS predicts the mean trailing edge wake but only LES offers a detailed view of the flow. Indeed, LESs identify flow phenomena in agreement with the experiment, with sound waves emitted from the trailing edge separation point that propagate upstream and interact with the lower blade suction side.

Viscous and potential forces on an advancing surface-piercing flat plate with a fixed drift angle

Transverse force and yaw moment acting on a surface-piercing flat plate with yaw angle and forward speed are studied. Following the Froude hypothesis, the problem is decomposed into a tail-separated forward flow and a bottom-tip-separated cross-flow parts. The objective is to quantify the role of different components in the resulted force and moment acting on the plate. A 3D potential flow code using distribution of Rankine sources and dipoles is used to solve the linear potential flow problem. The free-surface boundary condition is linearized based on the undisturbed flow velocity (Neumann–Kelvin linearization). The plate’s trailing-edge flow separation is modeled using a dipole distribution behind the plate on a linearized vortex sheet. The force and moment due to the cross-flow separation from the plate’s bottom tip are calculated by means of a slender body cross-flow method with a rigid free-surface boundary condition. Hence the free-surface effects are confined in the forward flow problem and neglected in the cross-flow problem. Simulations are carried out for different aspect ratios and drift angles. The plate’s thickness is taken into account. Convergence and sensitivity studies are performed carefully. Results are compared with previous experimental and numerical studies. The importance of the second-order forces and validity of the linear assumptions are touched upon. The overall agreement of the results is acceptable. The relative importance of the two viscous and potential components are studied. It is shown that the decomposed problem could follow the experimental results up to a relatively high Froude number.

Influence of self-adaptive hairy flaps on the stall delay of an airfoil in ramp-up motion

It is known in the case of some birds that the coverts on the upper side of their wings pop-up under critical flight conditions such as the landing approach, thus acting like a brake on the spread of flow separation. Taking experimental investigations as its basis, this paper deals with the influence of various configurations of self-adaptable hairy flaplets located on the lower half of the wing and with chord-length c (dense rows of slender elastomeric flaps, L=0.05c, 0.1c, 0.2c) on the flow around an NACA0020 airfoil at low Reynolds number flow (Re=77×103). Flow evolution along the airfoil when in ramp-up motion (α0=0, αs=20°, reduced frequency k=0.12) was measured with and without hairy flaps, with growth in the chord-normal thickness of the separation region above the airfoil investigated in order to determine stall onset time Ts. Whereas small flaps with L=0.05c do not change the overall stall process, it was possible to use configurations with L=0.1c (double-row, triple-row configuration) to delay stall onset Ts by a factor of around 2–4 when compared with the clean airfoil. The motion of the flaps and the flow field were measured simultaneously at high temporal resolution using high-speed PIV. Correlation between flap motion and velocity distribution showed that backflow induced by vortex structures is indeed prevented by the hairy flaps. A significant difference was identified in the shear-layer roll-up process, which was almost regular and locked with the fundamental frequency on the covered airfoil with no signs of non-linear growth over longer periods. By way of contrast, in the case of the clean airfoil the early merging of the shear-layer vortices and a rapid increase in the thickness of the separation region were observed. It is therefore concluded that mode locking is achieved between flap rows with an interspacing of 0.15c−0.2c, while the fundamental shear-layer roll-up wavelength measured (λ0≈0.15c−0.2c) indicates the relevance of flap row arrangement at the specific Reynolds number. Furthermore, interaction between shear-layer vortices and flaps in the row furthest downstream leads to the beneficial modification of the trailing edge flow in a way which increases bound circulation.

Effects of blunt trailing edge flow discharge in supersonic regime

Flow discharge at the blunt trailing edge of bodies immersed into supersonic stream alters significantly the downstream aerodynamic characteristics. This paper studies the changes in the base region topology for a wide range of purge ejection, using Unsteady Reynolds Averaged Navier–Stokes solutions. At low purge rate the base pressure increases, resulting in a reduction of the trailing edge shock wave intensity. Supersonic purge flows diminishes the base pressure and thus strengthens the shock wave, a linear relationship between the base pressure and the shock strength was obtained for the investigated conditions. Furthermore, an auxiliary shock wave system was observed to form upstream of the main fish tail shock at certain discharge rates.

On the estimation of wall pressure coherence using time-resolved tomographic PIV

Three-dimensional time-resolved velocity field measurements are obtained using a high-speed tomographic Particle Image Velocimetry (PIV) system on a fully developed flat plate turbulent boundary layer for the estimation of wall pressure fluctuations. The work focuses on the applicability of tomographic PIV to compute the coherence of pressure fluctuations, with attention to the estimation of the stream and spanwise coherence length. The latter is required for estimations of aeroacoustic noise radiation by boundary layers and trailing edge flows, but is also of interest for vibro-structural problems. The pressure field is obtained by solving the Poisson equation for incompressible flows, where the source terms are provided by time-resolved velocity field measurements. Measured 3D velocity data is compared to results obtained from planar PIV, and a Direct Numerical Simulation (DNS) at similar Reynolds number. An improved method for the estimation of the material based on a least squares estimator of the velocity derivative along a particle trajectory is proposed and applied. Computed surface pressure fluctuations are further verified by means of simultaneous measurements by a pinhole microphone and compared to the DNS results and a semi-empirical model available from literature. The correlation coefficient for the reconstructed pressure time series with respect to pinhole microphone measurements attains approximately 0.5 for the band-pass filtered signal over the range of frequencies resolved by the velocity field measurements. Scaled power spectra of the pressure at a single point compare favorably to the DNS results and those available from literature. Finally, the coherence of surface pressure fluctuations and the resulting span- and streamwise coherence lengths are estimated and compared to semi-empirical models and DNS results.