Leading-edge Separation Bubble Research Articles

Analysis of coupling supersonic turbine stage with rotating detonation combustor under different turbine parameters

The rotating detonation gas turbine boasts several advantageous characteristics, including self-pressurization, minimal entropy increase, high thermal efficiency, and a high thrust-to-weight ratio. These features make it a promising candidate for the next generation of aerospace power devices. This study investigates the flow characteristics and performance attributes of the supersonic turbine stage coupled with a rotating detonation chamber based on different rotational speeds, blade ratio of rotor and stator, and hub radius, and uses DMD (Dynamic Mode Decomposition) to analyze the main modes of the supersonic turbine. The research identifies distinct wave modes in both the stator and rotor blades of the supersonic turbine. In the stator blades, the modes include waveless contact modes, opposite side λ-wave oblique shock wave modes, and derived modes from opposite side λ-wave oblique shock waves. In the rotor blades, the modes encompass non-separation modes (waveless modes, asymmetric dual λ-wave modes, oblique cut wave modes, and single-side λ-wave modes on the pressure surface), leading-edge stator excitation single separation bubble modes, saddle-type separation modes, and dual separation bubble modes. The efficiency and power of the supersonic turbine stage increase with the rotational speed. The turbine efficiency is highest when the blade ratio is 10:7. Under multi-wave mode conditions, detonation waves demonstrate adaptive self-regulation abilities, ultimately achieving equilibrium in speed, detonation wave height, and spacing. The dominant modes in the stator blades and combustion chamber are governed by detonation waves and slip lines, while in rotor blades, the primary mode is the flow-directional vortex mode governed by detonation waves and slip lines at a frequency lower than their propagation frequency. This research holds certain reference significance for understanding the internal flow characteristics of rotating detonation gas turbines, as well as discussing the patterns of efficiency, power, and load characteristics, thereby offering valuable insights for the practical application of rotating detonation gas turbines.

Water tunnel testing of downwind yacht sails

Downwind yacht sails, such as spinnakers, are low-aspect-ratio highly cambered wings with a sharp leading edge. They are characterised by substantial three-dimensional flow separation and are thus modelled with difficulty with numerical simulations. Furthermore, accurate full-scale validation data are not available. The first quantitative flow measurements have only recently been achieved in water tunnels. In this study, we aim to provide guidelines on this emerging sail testing methodology. We consider six model-scale rigid models at average-chord-based Reynolds numbers ranging from 5 870 to 61 870. A critical Reynolds number is identified, below which relaminarisation of the reattached boundary layer downstream of the leading edge separation bubble occurs. Both lift and drag increase monotonically at subcritical Reynolds numbers while remaining about constant at transcritical Reynolds numbers. The critical Reynolds number decreases with increasing incidence and is insensitive to the blockage ratio. Spinnakers are normally sailed in front of the mainsail, whose circulation is found to generate an approximately 3∘\\documentclass[12pt]{minimal} \\usepackage{amsmath} \\usepackage{wasysym} \\usepackage{amsfonts} \\usepackage{amssymb} \\usepackage{amsbsy} \\usepackage{mathrsfs} \\usepackage{upgreek} \\setlength{\\oddsidemargin}{-69pt} \\begin{document}$$3^{\\circ }$$\\end{document} upwash on the spinnaker and higher flow velocity on both sides of it. These findings provide guidelines for the experimental testing of spinnaker-like wings in water tunnels and provide new insights into the flow and experimental testing of highly cambered wings with massive flow separation at low Reynolds numbers.

Dynamic Stall Control for a Vertical-Axis Wind Turbine Using Plasma Actuators

Dynamic stall is a major challenge for vertical-axis wind turbines (VAWTs) as it negatively impacts the aerodynamic performance of VAWTs, resulting in a reduction in net power output from the turbines. This study explores the use of alternating current dielectric barrier discharge plasma actuators for controlling dynamic stall on a VAWT. The flow control process is resolved using a two-dimensional unsteady Reynolds-averaged Navier–Stokes equations simulation with a Reynolds stress turbulence model adopted. The macro-effect of the plasma discharge on mean flow is modeled by an empirical body force model. The detailed flow actuation procedure is simulated. Numerical results show that the flow control process is dominated by the spanwise vortices, which result from the interaction between the plasma discharge and separated flow. The spanwise vortices not only prevent the development of a leading-edge separation bubble into a dynamic stall vortex, but also suppress the flow separation at the trailing edge. The application of plasma forcing has been found to greatly enhance the aerodynamic performance of the turbine in terms of the moment coefficient and averaged power coefficient of the turbine blades. Furthermore, we examine how the positioning of the plasma actuators and the freestream wind velocity affect flow control. The optimal actuation location is determined, and the plasma actuators are seen to be able to substantially enhance the net power output of the blades at all freestream speeds.

The effect of leading-edge shape on separation-induced transition on the suction surface of a controlled-diffusion airfoil

To investigate the effect of leading-edge shape on separation-induced transition on suction surface, a large eddy simulation is performed on two compressor controlled-diffusion airfoils: a conventional one with an elliptical leading edge and an optimized one with a curvature-continuous design based on the B-spline description. The Reynolds number based on inflow velocity and chord length is 4.5×105. The critical angle of attack +4°, over which the aerodynamic loss rises sharply, is chosen for simulation. Two transitions are observed on the suction surface, one near the leading edge and the other at 40% chord length. The primary difference between the two airfoils lies in the leading-edge transition, which also leads to the distinction of fluctuating velocity amplitude and energy loss in the subsequent development of boundary layer flow. In order to provide an insight into the transition mechanism, the frequency spectrum analysis is conducted, and the results indicate that the amplification of disturbances during transition is dominated by Kelvin–Helmholtz instability. The mechanisms of energy transport and dissipation are analyzed, and the influence of leading-edge curvature on the initial state of boundary layer flow is elucidated from a dynamic perspective. The results show that continuous and large curvature distributions are more conducive to suppressing the formation of leading-edge separation bubble and delaying the onset of transition.

Fluorescent oil-film applied to measure skin friction of flat plate with leading edge separation bubble in low Reynolds number

Under low Reynolds number conditions, the airfoil is prone to a laminar separation bubble (LSB), which causes problems such as lift drop and drag increase. In this work, the fluorescent oil film was applied to study the leading edge LSB from the perspective of skin friction. The experimental study is carried out on a flat plate with the leading edge LSB under the Reynolds number of [Formula: see text] (the characteristic length of the plate chord). The camera captures the leading edge LSB that is formed by airflow separation and reattachment, and the flow in the bubble is opposite to the free flow. The position of separation and reattachment is determined by the distribution of skin friction, which is consistent with the visual image of the oil film. Applying the Poincare–Bendixson (P-B) index formula to test the topology of the separation bubble skin friction, the quantitative relations between the critical points and the boundary switching points in the closed region of several permeable boundary all confirm this topology law. The results show that the fluorescent oil film can quantitatively map the flow structure of the LSB, and further verify the advantage of this method in the visual measurement of the complex flow structure.

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

Insight into the effects of leading edge delamination on the aerodynamic performance of an airfoil and wind turbine

Severe erosion or delamination at the leading edge of blades has an adverse influence on the power loss of wind turbines. Obtaining relevant quantitative data and its mechanisms would help in the efficient management and maintenance of these turbines. In the present work, detailed analysis on the effects of various levels of leading edge delamination on the aerodynamics and flow characteristics of an S809 airfoil and, in turn, on the performance of the NREL phase VI rotor is conducted. The results indicated that an elongated leading edge separation bubble and two enclosed vortex systems appeared in the delaminated region of the airfoil when the delamination levels aggravate to a certain degree, while the trailing edge separation vortex can be observed for all delaminated airfoils. The aerodynamic performance of the wind turbine decreases with increasing delamination depth and increases with growing delamination depth. The most influential delamination length is 1%c, and the maximum power loss is 23.09% for the blade with the severe delamination case. A small area of deep leading edge delamination or defects may cause great power loss, which should not be ignored in practice engineering.

Turbulent flow around circular arcs

The flow around a circular arc is governed by the effect of the sharp leading edge and the arc's curvature. There is a range of incidences where a leading-edge separation bubble (LESB) is formed on the convex side of the arc, and the reattached boundary layer separates further downstream. Akin to foils and cylinders, for increasing values of the Reynolds number, the boundary layer turns from laminar to turbulent resulting in a step change in the forces, here termed force crisis. This phenomenon is characterized experimentally for an arc with a camber-to-chord ratio of 0.22 and for a range of the Reynolds number from 53 530 to 218 000. Forces are measured both in a towing tank and in a water tunnel, and particle image velocimetry is undertaken in the water tunnel. In stark contrast to cylinders, where the force crisis is associated with the laminar-to-turbulent transition of the boundary layer, here, it is found to be associated with the suppressed relaminarization of the boundary layer. In fact, the LESB is always turbulent at the tested conditions, and relaminarization occurs up to a combination of critical angles of attack and critical Reynolds numbers. The critical angle of attack varies linearly with the Reynolds number. These results may contribute to the design of thin cambered wings, sails, and blades at a transitional Reynolds number such as the wings of micro aerial vehicles, swept wings in subsonic flight, turbomachinery blades, and the sails of autonomous sailing vessels.

Investigation of local unstart in a hypersonic scramjet intake at a Mach number of 6

Understanding the mechanism of local unstart in a mixed compression scramjet intake is vital for its stable off-design operation. In the present work, the local unstart behavior of a hypersonic scramjet intake is studied at a Mach number of 6, and a flight realistic Reynolds number of 8.4×106/m using a combination of experiments in a hypersonic wind tunnel and numerical computations. To study the effect of internal contraction ratio (ICR), experiments are performed on an intake model that is equipped with an all movable cowl in two modes: (a) cowl fixed at a particular ICR, and (b) cowl moved to a particular ICR during the steady test time of the wind tunnel. The intake exhibits started flow for ICR between 1.19 and 1.37, and locally unstarted flow for ICR ≥1.45. In the dynamic cowl experiments, the impingement of the forebody shock onto the inner cowl surface forms a leading-edge separation bubble, which moves upstream and detaches the cowl shock, leading to local unstart with subsonic spillage of mass near cowl lip. A shear layer emanating from the triple point of shock interaction divides this subsonic flow and the supersonic flow at the ramp side. The frequency content of the flow is different between the fixed and the dynamic cowl cases. Local unstart observed during fixed cowl experiments reveal low amplitude self-sustained oscillations around a frequency of 2.5 kHz. The dynamic mode decomposition (DMD) and proper orthogonal decomposition (POD) of numerical density contours reveal that the shear layer impingement on the cowl surface drives the self-sustained oscillations observed during local unstart.

The scales of the leading-edge separation bubble

The leading-edge separation bubble is a flow feature occurring on the suction side of thin foils as a result of separation at the sharp leading-edge followed by reattachment downstream along the chord. For a flat plate at zero angle of attack, the reattachment length of the bubble depends on the plate thickness t. At a nonzero incidence, instead, the underlying scale governing bubble length is not clear. To investigate, we undertake a critical review of experimental and theoretical studies and develop an analytical formulation to predict the reattachment length of plates both at zero and at small incidences. We focus on conditions where the bubble is turbulent, i.e., when transition occurs at a negligible distance from the point of separation. This occurs at thickness- and chord-based Reynolds numbers Ret≳104, Rec≳105. At angle of attack α = 0, we find that the reattachment length is xR≈4.8t when the chord-to-thickness ratio is c/t>12. At α>0, we find that xR/c=πσα2, where σ≈7.9 is the inverse of the growth rate of a turbulent free shear layer. These results allow estimating xR on the thin wings of, for example, aerial vehicles and yacht sails.

The influence of yaw on the unsteady surface pressures over a two-wheeled landing-gear model

Landing-gear noise is an increasing issue for transport aircraft. A key determinant of the phenomenon is the surface pressure field. Previous studies have described this field when the oncoming flow is perfectly aligned with the gear. In practice, there may be a cross-flow component; here its influence is investigated experimentally for a generic, two-wheel, landing-gear model. It is found that yaw angles as small as 5° cause significant changes in both overall flow topology and unsteady surface pressures. Most notably, on the outboard face of the leeward wheel, large-scale separation replaces predominantly attached flow behind a leading-edge separation bubble. The effect on unsteady surface pressures includes marked shifts in the content at frequencies in the audible range, implying that yaw is an important parameter for landing-gear noise, and that purely unyawed studies may not be fully representative of the problem.

Different noise generation mechanisms of a controlled diffusion aerofoil and their dependence on Mach number

The self-noise of an isolated controlled-diffusion aerofoil is investigated using direct noise computations. The high-order finite difference solver HiPSTAR is used to conduct four large eddy simulations and one direct numerical simulation. The motivation is to investigate the multiple sources of aerofoil self-noise on a realistic compressor blade geometry, the physical mechanisms leading to the generation of sound, and the influence of compressibility effects. The angle of attack is 8°, the chord-based Reynolds number is 105, and results obtained for four values of the free-stream Mach number [0.2, 0.3, 0.4, 0.5] are compared. For those flow parameters, the pressure side is fully laminar, whereas a separation bubble is present on the suction side close to the leading edge that promotes transition to turbulence. The size of the separation bubble is found to increase with the Mach number. Two noise sources are observed, one at the trailing edge and one in the leading edge transition/reattachment region. The first has a broadband, low frequency spectrum, while the second displays a tone whose frequency depends on the local Mach number. Because the leading edge separation bubble is very small, the associated tone frequency is high and requires a significantly finer grid to faithfully resolve the acoustic propagation than what is typically deemed sufficient. Finally, cross correlations between the surface pressure and the far-field pressure reveal that the pressure fluctuations reaching the trailing edge are initially generated in the transition/reattachment region, which indicates that the trailing edge noise is a consequence of the pressure fluctuations generated by the separation bubble.

High sensitive flexible hot-film sensor for measurement of unsteady boundary layer flow

The detailed flow characteristics in the boundary layer are critical for physical understanding of complex flow and aerodynamic performance optimization of aircraft. In this study, high sensitive micro-electro-mechanical system flexible hot-film sensor was developed for measurement of unsteady boundary layer flow. The effects of annealing temperature and time on the temperature coefficient of resistance (TCR) were studied, the highest TCR value of 5100 ppm °C−1 was obtained after 6 h of vacuum thermal annealing treatment at 400 °C. The flow transition and separation characteristics developed on two-dimensional NACA 0012 airfoil at various Reynolds numbers and angles of attack were investigated by using flexible hot-film sensors with validation by surface pressure distribution and aerodynamic force coefficient. The accurate locations of onset, 50% intermittent and end of transition were detected successfully. The behavior of spatial progression to the leading edge of transition point was found as the Reynolds number and angle of attack increases. The hot-film sensor measurements also demonstrated the static stall of the airfoil was caused by the leading edge separation bubble.

Turbulence Intensity Effects on a Leading-Edge Separation Bubble of Flat Plate Wing at Low-Reynolds Numbers

In this study, we experimentally investigate the effects of mainstream turbulence intensity (Ti) on a leading-edge separation bubble under low-Reynolds number (Rec) conditions range of 2.0 × 104 to 6.0 × 104. We used a flat plate to fix a separation point at the leading edge. Also, we visualized the behavior of the leading-edge separation bubble using the smoke wire technique and Particle Image Velocimetry (PIV) measurement. Furthermore, we measured the effect of Ti on the turbulent transition process in the separated shear layer using a hot-wire anemometer. The results indicate that the bypass transition for large Ti causes the turbulent transition, and so accelerates the reattachment of the separated shear layer. The results show that the bypass transition promotes the reattachment of the separated shear layer to maintain the leading-edge separation bubble on the upper surface even at high angles of attack, increasing the stall angle.

Effect of Leading-Edge Optimization on the Loss Characteristics in a Low-Pressure Turbine Linear Cascade

This paper presents a numerical study on the aerodynamics loss reduction characteristics after the leading-edge (LE) optimization in a low-pressure turbine linear cascade. The LE was optimized with a simple and practical method of “Class Function/Shape Function Transformation Technique” (CST). The simulation conditions, covering the whole working range, were independently determined by incidence, Reynolds number and Mach number. Quantitative loss analyses were carried out with a loss breakdown method based on volumetric integration of entropy production rates. To understand the reason of loss reduction, the local sources at different operating points were identified with entropy production rates. The results showed that LE optimization with the CST method played a positive role in decreasing the total losses, and the working range with lower loss was extended. The profile loss and the endwall loss were significantly reduced by the LE optimization, which were also verified to be the major causes of the total loss reduction by loss breakdown. The decrease of profile loss can be attributed to the boundary layer near the LE region and the boundary layer of downstream at off-design incidence. The reduction mostly came from the pressure side at negative incidence, while came from the suction side at the positive incidence. The endwall loss was decreased markedly about 2.5%–5% by the LE optimization at the incidence of‒12°, which was 1% at the incidence of 12°. The mechanism for the endwall loss reduction at different incidences was different from each other. At negative incidence, the LE optimization diminished the corner separation vortex on the pressure side. While at positive incidence, the benefits came from three aspects, i.e., reduced suction LE separation bubbles close to the endwall, reduced passage vortex strength, and weakened shear process between passage vortex and trailing shed vortex. The loss of the downstream zone was relatively lower than that of the profile losses and the endwall losses. The effect of LE optimization on the loss of the downstream zone at different conditions was complex and it depended both on the profile boundary layer behavior at the suction trailing edge and on the passage vortex strength.

Effects of leading edge defects on aerodynamic performance of the [formula omitted] airfoil

Defects at the leading edge of blades are an important source of power loss for wind turbines. In the present work, a systematic study is carried out on two typical types of defects, i.e., the surface concaved deep defects and surface distributed shallow defects on the S809 airfoil. Different defect shapes, ranges, equivalent depths and their influence mechanisms are investigated via CFD. For deep defects, an enclosed vortex is formed in the defect cavity, which suppresses the momentum exchange between the external flow and internal flow, so that the airfoil aerodynamic performance is highly sensitive to the defect opening range and is little affected by the defect shape and equivalent depth. Flow around the leading edge is strongly hindered by the deep defects and an elongated leading-edge separation bubble is formed at the suction side of the airfoil. At large angles of attack, flow at the suction side is dominated by flow separation at both the leading edge and trailing edge. In contrast, for shallow defects, all the defect equivalent depth, opening range and shape can significantly influence the airfoil aerodynamic performance; and quantitatively, the effect of defect equivalent depth is the most significant. For the present simulation cases with defects, the maximum lift coefficient is notably decreased (by 35% to 61%) accompanied by a sharp increase in drag coefficient (by 131% to 217%). Under dynamic pitching motions, the opening of the dynamic lift (drag)-coefficient hysteresis curve is effectively enlarged. The present work aims to provide an important reference for the maintenance and management of turbine blade defects.

The onset of dynamic stall at a high, transitional Reynolds number

Dynamic stall due to a ramp-type pitching motion is investigated on the NACA 0012 airfoil at chord Reynolds number of $Re_{c}=1.0\times 10^{6}$ through the use of wall-resolved large-eddy simulation. Emphasis is placed on the unsteady boundary-layer interactions that develop as the airfoil approaches stall. At this Reynolds number it is shown that turbulent separation moves upstream across much of the airfoil suction surface. When turbulent separation reaches the leading-edge separation bubble, a bursting event is initiated leading to a strong coherent leading-edge vortex structure. This vortex wraps up the turbulent shear layer to form a large dynamic stall vortex. The use of large-eddy simulation elucidates the roll of the laminar separation bubble in defining the onset of the dynamic stall process. Comparisons are made to identical simulations at lower Reynolds numbers of $Re_{c}=0.2\times 10^{6}$ and $0.5\times 10^{6}$. This comparison demonstrates trends in the boundary-layer mechanics that explain the sensitivity of the dynamic stall process to Reynolds number.

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.

Large Eddy Simulation of an asymmetric spinnaker

This work presents a numerical study into the aerodynamics of downwind sailing using different methods for modelling turbulence, comparing Large Eddy Simulation (LES) and Reynolds-Averaged Navier-Stokes (RANS) methods. Previously published experimental data of downwind sails have shown areas of flow separation and vortex structures at the leading and trailing edge of the gennaker, making these sails challenging to model accurately. The ability of LES to model the transition to turbulence within separated shear layers can lead to an accurate prediction of the leading edge separation bubble, which significantly influences the flow field around the top half of the sail, and downstream of the layer's reattachment. The transient nature of the LES solution allows the computation of the creation and shedding of unsteady vortices at the leading edge and downstream of the sail draft. The effect of the vortex rolls being convected towards the trailing edge is to generate a boundary layer which is more resistant to separation. Comparison with the pressure distribution measured in experiment shows that LES correctly predicts the location of separation, while RANS erroneously predicts a larger region of stalled flow which limits the suction on the sail. As a result, the overall drive and side forces computed by the LES are in better agreement with the experimental data, with less than 3.5% error, that the RANS which underestimates their magnitude by more than 14%. However, it should be noted that the accuracy of the LES calculations has not been demonstrated, and so the results should be treated with caution and not considered conclusive.