Leading-edge Laminar Separation Research Articles

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.

Characterization of Leading-edge Laminar Separation Bubble for various Angles of Incidence and Reynolds Numbers

Characterization of Leading-edge Laminar Separation Bubble for various Angles of Incidence and Reynolds Numbers

On the use of synthetic inflow turbulence for scale-resolving simulations of wetted and cavitating flows

The Delft Twist 11 Hydrofoil is a common test case for investigating the interaction between turbulence and cavitation modelling in computational fluid dynamics. Despite repeated investigations, results reported for the lift and drag coefficient are accompanied by significant uncertainties, both in experimental and numerical studies. When using scale-resolving approaches, it is known that turbulent fluctuations must be inserted into the domain in order to prevent the flow from remaining laminar around the body of interest, although this has been overlooked until now for the present test case. This work investigates the errors occurring when a laminar inflow is applied for mildly separated or attached flows, by employing the partially averaged Navier–Stokes equations with varying values for the ratio of modelled-to-total turbulence kinetic energy, and with varying grid densities. It is shown that depending on the grid resolution laminar leading edge separation can occur. When turbulent fluctuations are added to the inflow, the leading edge separation is suppressed completely, and the turbulent separation zone near the trailing edge reduces in size. The inflow turbulence has a large effect on the skin friction, which increases with increasing turbulence intensity to a limit determined by the grid resolution. In cavitating conditions the integral quantities are dominated by the shedding sheet cavity. The turbulence intensity has little effect on the pressure distribution, leading to a largely unaffected sheet cavitation, although the shedding behaviour is affected. It is shown that, especially in wetted flow conditions, with scale-resolving methods inflow turbulence is necessary to match the experimental flow field.

Numerical study of bluntness effects on laminar leading edge separation in hypersonic flow

Bluntness effects on laminar hypersonic leading edge separation are investigated numerically at Mach number $M\approx 10$, unit Reynolds number $Re=1.3\times 10^{6}~\text{m}^{-1}$, specific enthalpy $h_{o}=3.1~\text{MJ}~\text{kg}^{-1}$ and wall-to-stagnation temperature ratio $T_{w}/T_{o}=0.1$. Such effects are important from an experimental point of view and because bluntness can affect a separated flow favourably or adversely. In this study, two blunt leading edge cases of small radius ($15~\unicode[STIX]{x03BC}\text{m}$) and large radius ($100~\unicode[STIX]{x03BC}\text{m}$) are investigated. A comparison with the idealised sharp leading edge case is also given. General flow features and surface parameters such as the shear stress, pressure and heat flux are presented and analysed. The results have also been interpreted in terms of Cheng’s displacement-bluntness similitude parameter affecting the size of separation. Previous experiments by Holden delineated small and large bluntness effects based on Cheng’s parameter and considering small to moderate separated regions. In this study, leading edge separation was found to suppress the favourable effect of bluntness in delaying separation. Bluntness, furthermore, seemed to promote the appearance of secondary vortices within a main separated region. Analysis of reverse flow boundary layer profiles such as velocity, pressure and temperature is also given. It is shown that bluntness accentuates large transverse gradients. This in turn adversely effects the reverse flow boundary layer leading to the appearance of secondary vortices.

Unsteady Flow Physics of Airfoil Dynamic Stall

A series of wind-tunnel experiments was conducted on a dynamically pitching NACA 0012 airfoil in order to understand the unsteady flow physics associated with dynamic stall at a transitional Reynolds number of . The airfoil was pitched about the quarter-chord axis following a linear ramp maneuver at . A series of high-frequency unsteady surface pressure measurements were acquired in order to investigate the time-dependent behavior of the flow in the immediate vicinity of the airfoil. These surface pressure measurements actively displayed the evolution of the airfoil boundary layer, leading-edge laminar separation bubble, and dynamic stall vortex. A detailed time-frequency analysis of surface pressure measurements was performed in order to identify the dominant frequencies associated with the unsteady laminar separation bubble. This analysis was also used to study the evolution of the length scales associated with flow structures near the airfoil leading edge, which emerged before the formation of the dynamic stall vortex. Finally, time-resolved particle image velocimetry measurements were acquired to study the evolution of the off-body flowfield during the development of dynamic stall.

Unsteady aerodynamic effects in small-amplitude pitch oscillations of an airfoil

High-fidelity wall-resolved large-eddy simulations (LES) are utilized to investigate the flow-physics of small-amplitude pitch oscillations of an airfoil at Rec=100,000. The investigation of the unsteady phenomenon is done in the context of natural laminar flow airfoils, which can display sensitive dependence of the aerodynamic forces on the angle of attack in certain “off-design” conditions. The dynamic range of the pitch oscillations is chosen to be in this sensitive region. Large variations of the transition point on the suction-side of the airfoil are observed throughout the pitch cycle resulting in a dynamically rich flow response. Changes in the stability characteristics of a leading-edge laminar separation bubble has a dominating influence on the boundary layer dynamics and causes an abrupt change in the transition location over the airfoil. The LES procedure is based on a relaxation-term which models the dissipation of the smallest unresolved scales. The validation of the procedure is provided for channel flows and for a stationary wing at Rec=400,000.

Numerical Simulations of Turbulent Flow over Airfoils Near and During Static Stall

Reynolds-averaged Navier–Stokes and hybrid large-eddy/Reynolds-averaged Navier–Stokes simulations of turbulent flow past an Aérospatiale A-Airfoil near stall at , , and a NACA 0012 airfoil under static stall conditions (, , ) are described in this paper. In the flow past the A-Airfoil, comparisons with surface skin-friction coefficient and pressure coefficient distribution are generally in good agreement with experimental measurements. Comparisons with experimental velocity profile data and Reynolds-stress data are also generally favorable. Leading-edge laminar separation and turbulent reattachment is predicted when the Menter–Langtry correlation-based transition model is used in combination with either Reynolds-averaged Navier–Stokes or large-eddy/Reynolds-averaged Navier–Stokes strategies, but the level of trailing-edge separation is underpredicted, relative to experimental data and to results obtained without the inclusion of transition model. For the case of flow past a NACA 0012 airfoil under static stall conditions, results from the large-eddy/Reynolds-averaged Navier–Stokes simulations exhibit a sensitivity to mesh refinement, with finer spanwise mesh resolution leading to light stall, characterized primarily by trailing-edge separation, and coarser spanwise mesh resolution leading to deep stall, characterized by the presence of a stabilized leading-edge vortex. Calculations that include wind-tunnel wall effects show better general agreement with experimental measurements but also display the aforementioned sensitivity to spanwise mesh refinement.

Modelling roughness effects for transitional low Reynolds number aerofoil flows

Roughness modelling at low Reynolds numbers of O(104–105) is of practical importance for micro air vehicles. This paper investigates the roughness modelling behaviour of the low Reynolds number shear stress transport model and the γ-Reθ shear stress transport model. Both include modelling flow transition and surface roughness effects. The roughness effects are modelled as sand grain roughness. A series of simulations using the two models have been performed on a NACA0012 aerofoil with comparisons to available experimental data. The results show that both of the models have the capability to reasonably predict the leading edge laminar separation bubble, transition and skin friction and, therefore, lift and drag on smooth surfaces. However, the two models behave very differently for the rough surface aerofoil. While the low Reynolds number shear stress transport model performs well, the γ-Reθ model fails to predict the transition on the rough aerofoil surface, resulting in inaccurate lift and drag prediction.

Dynamic Roughness as a Means of Leading-Edge Separation Flow Control

The aircraft industry, as a whole, has been deeply concerned with improving the aerodynamic efficiency of current and future flight vehicles, particularly in the commercial and military markets. However, of particular interest to the field of aerodynamics is the elusive concept of a workable flow control mechanism. Effective flow control is a concept which if properly applied can increase aerodynamic efficiency. Various concepts and ideas to obtain successful flow control have been studied in an attempt to reap these rewards. Some examples include boundary layer blowing (steady and periodic), suction, and synthetic jets. The overall goal of flow control is to increase performance. The specific objectives of flow control include: 1) delay or eliminate flow separation, 2) delay boundary layer transition or 3) reduce skin friction drag. The purpose of this research is to investigate dynamic surface roughness as a novel method of flow control technology for external boundary layer flows. As opposed to standard surface roughness, dynamic roughness incorporates small time dependent perturbations to the surface of the airfoil. These surface perturbations are actual humps and/or ridges that are on the scale of the laminar boundary layer, and oscillate with an unsteady motion. Research has shown that this can provide a means to modify the instantaneous and mean velocity profile near the wall and favorably control the existing state of the boundary layer. The results of this study have shown that dynamic roughness can be a viable alternative in delaying and/or eliminating the leading edge laminar separation bubble and hence reaping some of the rewards of an effective flow control system, while also maintaining some physical advantages over other techniques.

On the coherent structures and stability properties of a leading-edge separated aerofoil with turbulent recirculation

AbstractThe present study is motivated by a need to produce stability modes to assist in the understanding and control of unsteady separated flows. The flow configuration is a NACA 0015 aerofoil with laminar leading-edge separation and turbulent recirculation. In previous water tunnel experiments, this flow configuration was measured in an unperturbed (uncontrolled) separated state, and a harmonically perturbed (controlled) reattached state. This study presents numerical data of the unperturbed case, and recovers stability modes to describe the evolution of perturbations in this environment. The unperturbed flow is numerically generated using large eddy simulation. Its temporal properties are quantified via a Fourier analysis of the velocity time history at selected points in space. The leading-edge shear layer instability is characterized by instantaneous vortex structures, and the bluff body shedding is illustrated by proper orthogonal decomposition modes. Statistical measures of the velocity field agree well with the water tunnel measurements. Finally a stability analysis is undertaken using a triple decomposition to distinguish between the time averaged field, the unsteady scales of motion, and a coherent wave (perturbation). This analysis identifies that perturbations in the region immediately downstream of the separated shear layer have the highest spatial growth rates. The associated frequency is of the order of the sub-harmonic of the shear layer instability.

Full-scale pressure measurements on a Sparkman and Stephens 24-foot sailing yacht

The aerodynamics of a Sparkman and Stephens 24-foot sailing yacht was investigated. Full-scale pressure measurements were performed on the mainsail and the genoa in upwind condition. Pressure taps were adopted to measure the pressures on three horizontal sections on the windward and leeward sides of the two sails. Several trims and apparent wind angles were tested. The present paper shows the pressure distributions on the sails and correlates the measured pressures with the flow pattern. In particular, leading-edge laminar separation bubble, turbulent reattachment and turbulent separation are discussed. Pressure measurements are also adopted to draw some trim guidelines.

Bat-Inspired Wing Aerodynamics and Optimization

A S ENGINEERS search to make micro air vehicles more stable, maneuverable, and efficient, many are turning toward biology for inspiration. Bats adapt to their environment by morphing their wings in flight. Depending on their niche, certain species soar and glide [1], whereas others perform barrel rolls in nature [2] and can pull up to 4.5g in obstacle courses [3].Morphological changes afford bats great agility at low speeds, and they are able tomaintain stability and control at low Reynolds numbers, in which viscous effects and leading edge laminar separation bubbles cause nonlinearities in lift [4–6]. Bats achieve these feats with fingerlike jointed bone structures and flexible wing membranes. These unique traits allow them to change camber and twist in flight, unlike avian span changes. Recent studies investigate replicating flapping bat flight [7–9], whereas others focus specifically on flexible, membrane wing benefits at low Reynolds numbers for improvements in gust alleviation and delayed stall characteristics [10–12]. These wings are passive elements, and it is difficult to attach control surfaces to them for flight authority. By actively controlling and morphing flexible wings, conventional controllers can be replaced. Aircraft morphing allows single vehicles to have multiple functions, ideally with continuous lifting surfaces to alleviate drag and vibration and increase efficiency. This is the focus of Garcia et al. [13], which proposes a nonflapping wing with twist capabilities. Morphing enables tailoring of wing shapes to multiple flight regimes, from takeoff through cruise to landing [14–16]. Many mechanisms have been proposed for morphing, such as the smart joint, an active rigidity composite suited to actuating a batlike membrane wing in flight [17]. This low-profile device can be embedded at joints in the fingerlike skeletal wing structure as a bimorph actuator [18]. Whereas this work focuses on static wing configurations rather than morphing behavior, results help specify actuator requirements used for a variable camber and twist wing. In this work, two key features of bat flight are studied, uniquely evolved bat wing planforms adapted to environments, and capabilities afforded through variable camber and twist, and are applied to rigid, fixed wing designs for small man-made craft. The goal of this study is not to mimic natural bat flight, but to understand how certain aspects of bat flight apply to the engineering problem of wing design for micro air vehicles. II. Background and Motivation

Computational and experimental investigation of flow over a transient pitching hydrofoil

The present study is developed within the framework of marine structure design operating in transient regimes. It deals with an experimental and numerical investigation of the time–space distribution of the wall-pressure field on a NACA66 hydrofoil undergoing a transient up-and-down pitching motion from 0° to 15° at four pitching velocities and a Reynolds number Re = 0.75 × 10 6. The experimental investigation is performed using an array of wall-pressure transducers located on the suction side and by means of time–frequency analysis and Empirical Modal Decomposition method. The numerical study is conducted for the same flow conditions. It is based on a 2D RANS code including mesh reconstruction and an ALE formulation in order to take into account the foil rotation and the tunnel walls. Due to the moderate Reynolds number, a laminar to turbulent transition model was also activated. For the operating flow conditions of the study, experimental and numerical flow analysis revealed that the flow experiences complex boundary layer events as leading-edge laminar separation bubble, laminar to turbulent transition, trailing-edge separation and flow detachment at stall. Although the flow is relatively complex, the calculated wall pressure shows a quite good agreement with the experiment provided that the mesh resolution and the temporal discretization are carefully selected depending on the pitching velocity. It is particularly shown that the general trend of the wall pressure (low frequency) is rather well predicted for the four pitching velocities with for instance a net inflection of the wall pressure when transition occurs. The inflection zone is reduced as the pitching velocity increases and tends to disappear for the highest pitching velocity. Conversely, high frequency wall-pressure fluctuations observed experimentally are not captured by the RANS model. Based on the good agreement with experiment, the model is then used to investigate the effects of the pitching velocity on boundary layer events and on hydrodynamic loadings. It is shown that increasing the pitching velocity tends to delay the laminar-to-turbulence transition and even to suppress it for the highest pitching velocity during the pitch-up motion. It induces also an increase of the stall angle (compared to quasi-static one) and an increase of the hysteresis effect during pitch-down motion resulting to a significant increase of the hydrodynamic loading.

Aerodynamic Characteristics of Elliptic Airfoils in High Reynolds Numbers

In the present study, the aerodynamic characteristics of elliptic airfoils were numerically investigated based on the Reynolds-averaged Navier-Stokes equations and the Spalart-Allmaras one-equation turbulence model. Calculations were performed for several Reynolds numbers between 2.48 x 10 5 and 3.91x10 6 at freestream Mach numbers ranging from subsonic to transonic. The formation of the leading-edge laminar separation bubble and the flow separation characteristics at the trailing edge were studied, including the effects of Reynolds number and compressibility. Elliptic airfoils with different thicknesses were considered for the investigation of the thickness effect. The results were compared with those of NACA series airfoils with identical thicknesses.

A parametric study of LES on laminar-turbulent transitional flows past an airfoil

Low-Reynolds-number aerodynamic performance of small-sized air vehicles is an area of increasing interest. In this study, low-Reynolds-number flows past an SD7003 airfoil are investigated to understand important viscous features of laminar separation and transitional flow followed by the complicated behavior of the flow reattachment process. In order to satisfy the three-dimensional (3D) requirement of the code, a simple “3D wing” is constructed from a two-dimensional (2D) airfoil. A parametric study of large eddy simulation (LES) on the airfoil flows at Re = 60,000 is performed. Effects of grid resolution and sub-grid scale (SGS) models are investigated. Although 3D effects cannot be accurately captured owing to the limitation of the grid resolution in the spanwise direction, the preliminary LES calculations do reveal some important flow characteristics such as leading-edge laminar separation and vortex shedding from the primary laminar separation bubble on the low-Reynolds-number airfoil.

Turbulence Amplification with Incidence at the Leading Edge of a Compressor Cascade

Detailed measurements, with a two-component laser-Doppler velocimeter and a thermal anemometer were made near the suction surface leading edge of controlled-diffusion airfoils in cascade. The Reynolds number was near 700,000, Mach number equal to 0.25, and freestream turbulence was at 1.5% ahead of the cascade.It was found that there was a localized region of high turbulence near the suction surface leading edge at high incidence. This turbulence amplification is thought to be due to the interaction of the free-shear layer with the freestream inlet turbulence. The presence of the local high turbulence affects the development of the short laminar separation bubble that forms very near the suction side leading edge of these blades. Calculations indicate that the local high levels of turbulence can cause rapid transition in the laminar bubble allowing it to reattach as a short “non-burst” type.The high turbulence, which can reach point values greater than 25% at high incidence, is the reason that leading edge laminar separation bubbles can reattach in the high pressure gradient regions near the leading edge. Two variations for inlet turbulence intensity were measured for this cascade. The first is the variation ofmaximum inlet turbulence with respect to inlet-flow angle; and the second is the variation of leading edge turbulence with respect to upstream distance from the leading edge of the blades.

Computational prediction of flow around highly loaded compressor-cascade blades with non-linear eddy-viscosity models

Abstract A computational study is presented which examines the predictive performance of two variants of a cubic low-Re eddy-viscosity model when applied to the flow around two highly loaded compressor-cascade blades. Particular challenges are posed by the transitional nature of the flow and the presence of laminar leading-edge separation in off-design conditions. Comparisons are presented for local as well as integral flow parameters, including turbulence intensity and losses. The study demonstrates that the cubic model is able to predict, in accord with experimental data and in contrast to a linear base-line model, the influential leading-edge separation preceding transition, due to the suppression of turbulence-energy generation in the impinging zone ahead of the leading edge. This attribute and a greater sensitivity to streamline curvature enable the model to give, in some flow conditions, a more realistic description of the development of the boundary layer after transition and a better prediction of the loss at off-design incidence. However, the model also exhibits predictive weaknesses, among them an inappropriate delay in reattachment and transition to fully developed turbulence, resulting in insufficient sensitivity to adverse pressure gradient once transition has occurred.

Competing Mechanisms of Compressible Dynamic Stall

Earlier experiments have documented the onset of compressible dynamic stall either from the bursting of a leading-edge laminar separation bubble or from a leading-edge shock, depending on the Reynolds number and Mach number. For certain combinations of conditions, the supersonic flow and the bubble dynamics compete with each other. The consequent complex interactions lead to a newly discovered mechanism of dynamic stall onset. Details of these various mechanisms are discussed

Inlet turbulence distortion and viscous flow development in a controlled-diffusion compressor cascade at very high incidence

Detailed two-component laser Doppler velocimeter (LDV) measurements of the flow through a controlleddiffusion (CD) compressor cascade at a Reynolds number of about 700,000 and at a low Mach number are reported in this article. A very high-incidence angle (8 deg above design) was considered throughout this investigation, which included the full experimental characterization of the turbulence Held. The LDV measurements were fully automated and were all taken in coincidence mode, thus turbulent flow correlations could be determined. Most significant was the measurement of the distortion of the inlet freestream turbulence upstream of the blade leading edges. Such information is important in assessing viscous codes which incorporate transport equations to describe the turbulence within the flowfield. The laminar leading-edge separation bubble, which reattached turbulent, was enlarged on the suction surface of the blade. Consistent with measurements at lower incidence angles, the suction surface boundary layer remained attached over the rear part of the blade. The pressure side boundary layer initially showed little or no growth, however, it finally developed into a profile similar to a wall jet. The wake profiles showed significant asymmetry due to the high loading on the blades at the increased incidence angle.

On the abrupt turbulent reattachment downstream of leading-edge laminar separation

A theoretical investigation of unsteady marginal separations, in a classical unsteady boundary layer, is described. This suggests that the nonlinear properties of such separations may be largely responsible, at least initially, for the central features often observed experimentally in the abrupt turbulent reattachment of eddies that follow laminar separation near the leading edges of aerofoils. The theory reduces the local flow problem to the solution of a nonlinear integro-partial differential equation for the unknown scaled skin friction, which is also proportional to the decrement in displacement locally. Numerical solutions followed by linear and nonlinear analysis show that all short-length small disturbances are unstable whenever reversed flow occurs; but for sufficiently confined initial conditions the instabilities accumulate in such a way that a breakdown is forced after a finite time. In the nonlinear breakdown a double structure arises near the point of collapse, with an effective ‘shock’ taking the flow abruptly from the reversed sense just upstream to the forward sense just downstream. This predicted shock forms the suggested possible link with the sudden transition and turbulent reattachment of laminar eddies in practice, at rounded leading edges. Discussions of the turbulence effects, of the related phenomenon of dynamic stall, of other instabilities present, and of the theoretical repercussions of the work are presented also.