Laminar Separation Bubble Research Articles

Temperature and skin-friction maps on a lifting hydrofoil in a propeller wake

There are at least two direct links between the friction acting on the surface of a (slightly warmer or colder) body under the influence of an incompressible flow and the temperature distribution on the surface of the body itself. The first relies on the energy equation, which connects the evolution of the temperature distribution at the wall to the action of the skin-friction field. On the other hand, the equation of passive transport of temperature perturbations at the wall unveils a direct relationship between the celerity of propagation of thermal blobs and the friction velocity. Grounding on these relationships, this paper reports about the application of different methodologies, developed to determine skin-friction fields from surface temperature maps, to the analysis of the complex flow evolution on a lifting NACA 0015 hydrofoil immersed in a non-uniform, unsteady wake induced by a marine propeller. The adopted temperature-sensitive paint technique allows obtaining the global temperature data with the appropriate, high resolution in space and time. The approach based on the energy equation leads to an unconventional, single-snapshots optical-flow-like methodology, which allows obtaining time-resolved, relative skin-friction fields. The two approaches based on the displacement of the wall temperature perturbations enable the determination of time- and phase-averaged, quantitative skin-friction fields. One approach grounds on a classical tracking of the thermal disturbances via optical flow, the other one relies on minimizing the discrepancy between the celerity of propagation of thermal fluctuations and the behavior expected according to the Taylor hypothesis. The analysis of the skin-friction fields obtained via the three uncorrelated, complementary approaches discloses the evolution and the mutual interaction of different flow structures (manifolds) over the hydrofoil surface at lifting and zero-lift conditions: the hub and tip vortices, the blade wake, the laminar separation bubble, and the separation at trailing edge.

Effect of turbulence intensity and gradient of turbulence intensity on airfoil aerodynamic characteristics at low Reynolds number

Based on the complex flow field of vertical takeoff and landing (VTOL) aircraft with distributed propulsion, the influence of the turbulence intensity and gradient of turbulence intensity on the aerodynamic characteristics of two-dimensional airfoil under low Reynolds number was studied by solving the unsteady Reynolds averaged Navier-Stokes (URANS) Equation based on the c-type structural mesh and γ-Reθt transition model. The aerodynamic characteristics of NACA0012 airfoil at different turbulence intensities and Reynolds numbers are simulated and compared with the experimental data, which verifies the reliability of the low Reynolds number calculation method. Meanwhile, the effects of the different low Reynolds number and gradient of turbulence intensity on the aero-dynamic characteristics of airfoil are studied, and the effect mechanism of the turbulence on the flow field around airfoil is analyzed. It shows that the flow characteristics of the airfoil with high turbulence or Reynolds number are more stable, the separation bubble size is smaller, the flow separation is delayed, and the stall angle of attack is larger, but the effect of the two mechanisms on the earlier transition is different. The influence of the turbulence gradient on the airfoil is limited by the Reynolds number, and the flow separation, transition and reattachment of the airfoil with high turbulence gradient are advance. The generation and evolution of the laminar separation bubble are closely related to the turbulence intensity and Reynolds number, and its scale and location also affect the aerodynamic characteristics of the airfoil.

Dynamic Stall over a Pitching Natural-Laminar-Flow Airfoil

High-fidelity large-eddy simulations are performed to study the unsteady boundary layer sensitivity of a natural-laminar-flow airfoil (NLF-0414) undergoing dynamic stall in comparison to two traditional airfoil sections (NACA 0012 and SD 7003). In this work, the natural-laminar-flow airfoil at and pitches with a constant nondimensional rate of from an initial incidence angle of 4 deg to angles of attack beyond the onset of dynamic stall. It is found that the transition of the NLF-0414’s unsteady boundary layer during the pitchup is much more abrupt than the standard airfoil sections before the formation of the laminar separation bubble. Once the laminar separation bubble is established near the leading edge, the overall dynamic stall process of the NLF-0414 evolves very similarly to that of the SD 7003 and the NACA 0012. Minor differences are observed in the formation of the turbulent separation vortex near the trailing edge and convection of the dynamic stall and shear-layer vortices. The dynamic stall of the NLF-0414 airfoil more closely resembles that of the SD 7003 and the NACA 0012 at reduced Reynolds numbers, suggesting that the extended laminar flow acts to delay Reynolds number effects of traditional airfoil profiles.

Stability-Based Transition Transport Modeling for Unstructured Computational Fluid Dynamics at Transonic Conditions

In the present work the extension of a recently published transition modeling approach is proposed. In the modified variant of the approach, compressibility effects up to an onflow Mach number of 1 are taken into account using the compressible formulation of the Arnal–Habiballah–Delcourt transition criterion aiming at flow conditions and phenomena being typical for transport-type aircraft. Moreover, the Gleyzes criterion for the estimation of the transition onset location inside laminar separation bubbles was incorporated into the approach providing a higher accuracy for separation-induced transition points. The extensions were validated for the following test cases: The flat plate, the CAST10-2 airfoil, the NLR 7301 airfoil, and the Eppler 387 airfoil including grid resolution studies. Moreover, details with regard to the implementation and application of the extensions are given.

Investigation of laminar separation bubble over a supercritical airfoil in an incompressible flow

Supercritical airfoils have an unknown behavior at incompressible flow regime and Reynolds numbers lower than those related to their design point at transonic condition. In this work, boundary layer transition is studied over a supercritical airfoil by means of hot-film and pressure measurements completed with numerical simulations. The experiments are performed at chord-based Reynolds number of [Formula: see text]and Mach number of [Formula: see text] at different angles of attack. Hot-film measurement over the upper surface of the supercritical airfoil is carried out and the transition points are computed using the standard deviation of the signals. The upper surface pressure is also recorded and a peak in its second derivative is presented as the transition point generated by the laminar separation bubble mechanism. Moreover, an appropriate time-frequency analysis is applied to the hot-film signals to get an insight into the spectral content and development of the transitional boundary layer structures. On the other hand, two numerical codes are employed and the transition points obtained from numerical simulations are compared with the experimental outcomes. Results express a rapid change of the bubble position over the upper surface, as the angle of attack is increased to the value of [Formula: see text]. Laminar separation bubble is observed in the surface pressure distribution data and is well identified using its second derivative along the streamwise direction. The spectral characteristics of the boundary layer are satisfactorily explored including the streamwise fluctuations within the laminar flow, intermittent behavior of the transitional zone and the wide range of the spectrum in turbulent flow, thanks to the time-frequency analysis. A promising agreement is observed between the transition points computed by both the numerical and experimental studies and confirms the accuracy of findings achieved by the second derivative of surface pressure data, hot-film measurements and the reliability of the employed numerical transition models for optimization studies.

Development of an intermittency transport equation for modeling bypass, natural and separation-induced transition

The objective of this paper is to propose a new intermittency transport equation that is naturally capable of capturing the effects of free-stream turbulence and pressure gradient on bypass and natural transition without need for any extra parameters/terms to take into account the free-stream turbulence and the pressure gradient. This new intermittency transport equation is obtained by derivation and, in it, only its production term is modeled via two empirical functions in order to detect the onset location of transition and to control the growth rate of transition process. Only one sensing parameter is used to detect the transition onset for both natural and bypass transition. For the purpose of not altering the original form of the base turbulence model, the effective intermittency factor is used to describe the separation-induced transition. The base turbulence model remains unmodified in conjunction with the effective intermittency factor as a regulator to control the net turbulent generation rate. For the evaluation of model performance, the modeled intermittency equation is tested against (1) the transitional boundary layer on a flat plate with zero and non-zero pressure gradients, (2) the transitional flow over a compressor blade with a laminar separation bubble, and (3) the transitional flow over a wind turbine airfoil at various angles of attack. The present results are also compared to those of the transition model of Menter et al. [1] and those of the transition model of Langtry and Menter [2]. The evaluation results reveal that the new intermittency transport equation is capable of predicting the bypass, natural and separation-induced transition.

Low-Noise Synthetic Turbulence Tailored to Lateral Periodic Boundary Conditions

The present work is dedicated to turbulence synthesis tailored to lateral periodic boundary conditions for direct noise computations through compressible large eddy simulations. Synthetic turbulence can be essential for aeroacoustic applications when computing airfoil turbulent inflow noise or for accurately capturing the behavior of boundary layers. This behavior determines both trailing edge noise and complex flow structures such as laminar separation bubbles. For airfoil simulation purposes, spanwise periodic boundary conditions are usually considered. If synthetic perturbations are injected without observing the periodicity rule, strong spurious pressure waves are emitted and pollute the entire computational domain. In this work, the random Fourier modes method for turbulence generation is adapted in order to respect the spanwise periodicity constraint right at the computational domain inlet. This approach does not affect the turbulence properties such as the spectral shape and the turbulent kinetic energy decay. Since the emphasis is put on the generation and convection of the turbulence, only the turbulence convection region between the inlet and the airfoil is considered in this paper, without the airfoil. Two geometrical configurations are tested: the first one is a simple box with a constant mesh size, and the second one concentrates the fine cells on the area in front of the airfoil. In the second configuration, the computational cost is reduced by up to 25%, but more spurious noise is present because of interpolation areas between different grids using the Chimera method. Finally, the results’ reproducibility is assessed using different turbulence realizations.

Investigation on flow structure and aerodynamic characteristics over an airfoil at low Reynolds number—A review

This Review summarizes the progress in research on the flow structure and aerodynamic characteristics of an airfoil at a low Reynolds number encountered by near-space low-speed aircrafts and micro-air vehicles. The structures of several kinds of laminar separation bubbles and their effect are discussed by drawing on experimental and numerical results reported in the past few decades. The transition process in the separation bubble is detailed from various perspectives, including the receptive, primary instability, secondary instability, and break-down stage. The process of evolution of a coherent vortex structure that may affect the transition is discussed by analyzing the vortex dynamics in the separation bubble. Combined with the flow characteristics at a low Reynolds number and data on the airfoil, aerodynamic characteristics of the airfoil, such as the nonlinear effect and static hysteresis effect at a low angle of attack, are discussed.

Stability characteristics of different aerofoil flows at [formula omitted] and the implications for aerofoil self-noise

Three compressible-flow direct numerical simulations (DNS) of aerofoils were conducted at a chord based Reynolds number of Rec=150,000: a Controlled-Diffusion (CD) aerofoil at two incidences and an untripped NACA 6512-63 aerofoil at 0∘ angle of attack (AoA). The aim of the simulations was to help improve the understanding of the generation mechanisms of aerofoil self-noise. To this end, the time-averaged flows around the aerofoils were investigated using a multi-dimensional global stability analysis based on the response to very low-amplitude hydrodynamic initial pulses. For the CD aerofoil at 5∘ AoA and the untripped NACA 6512-63 aerfoil at 0∘ AoA, the base flows are found to be globally unstable and an acoustic feedback loop occurs regardless of the perturbation locations. For the CD aerofoil at a slightly higher incidence (8∘ AoA) where a laminar separation bubble appears at the leading edge and a turbulent attached boundary layer is present at the trailing edge, the initial perturbations decay with time for all initial pulse locations indicating that the base flow is globally stable. The results confirm that for cases where a laminar separation bubble is present on the aerofoil suction side at the trailing edge, independent of whether tones in the DNS exist or not, an unstable feedback loop exists. The laminar separation bubble at the aerofoil trailing edge on the suction side seems to be the only required prerequisite for triggering the acoustic feedback loop.

Unmanned Aircraft Systems Performance in a Climate-Controlled Laboratory

Despite many research studies focus on strategies to improve autopilot capabilities and bring artificial intelligence onboard Unmanned Aircraft Systems (UAS), there are still few experimental activities related to these vehicle performance under unconventional weather conditions. Air temperature and altitudes directly affect thrust and power coefficients of small scale propeller for UAS applications. Reynolds numbers are usually within the range 10,000 to 100,000 and important aerodynamic effects, such as the laminar separation bubbles, occur with a negative impact on propulsion performance. The development of autonomous UAS platforms to reduce pilot work-load and allow Beyond Visual Line of Sight (BVLOS) operations requires experimental data to validate capabilities of these innovative vehicles. High quality data are needed for a deep understanding of limitations and opportunities of UAS under unconventional flight conditions. The primary objective of this article is to present the characterization of a propeller and a quadrotor capabilities in a pressure-climate-controlled chamber. Mechanical and electrical data are measured with a dedicated test setup over a wide range of temperatures and altitudes. Test results are presented in terms of thrust and power coefficient trends. The experimental data shows low Reynolds numbers are responsible for degraded thrust performance. Moreover, details on brushless motor capabilities are also discussed considering different temperature and pressure conditions. The experimental data collected in the test campaign will be leveraged to improve UAS design, propulsion system modelling as well as to provide guidelines for safe UAS operations in extreme environments.

Laminar Separation Bubble Development on a Finite Wing

This experimental study considers the development of a laminar separation bubble formed over a finite wing and contrasts it with the flow evolution over a two-dimensional airfoil. The experiments were performed on a reference NACA 0018 airfoil model and a finite wing with an aspect ratio of 2.5 at a Reynolds number of 125,000 involving surface pressure and particle image velocimetry measurements. For equivalent effective angles of attack, the results reveal significant differences between the two- and three-dimensional configurations. Specifically, increasing the effective angle of attack causes the separation bubble to shift upstream on the two-dimensional airfoil, whereas its mean position and streamwise extent remain invariant to spanwise changes in the effective angle on the finite wing. Similarly, the dominant shear layer frequency and characteristics of shear layer rollers do not vary appreciably over the wingspan. This suggests that the sectional analogy between the local effective angle on the wing and that on the airfoil is not universally applicable to the wing sections subjected to the formation of laminar separation bubbles. Instead, the spanwise characteristics of the bubble on the wing are well approximated by those obtained on the airfoil at the angle of attack matching the effective angle of the wing root. It should be noted, however, that these findings are exclusive of the regions in the immediate vicinity of the wing tip and/or root, which were not considered in the present investigations.

Inspection of structures interaction in laminar separation bubbles with extended proper orthogonal decomposition applied to multi-plane particle image velocimetry data

This work reports the application of an extended proper orthogonal decomposition (E-POD) procedure to multi-plane particle image velocimetry (PIV) measurements describing the evolution of laminar separation bubbles (LSBs). Measurements were performed over a flat plate installed between adjustable end-walls providing a prescribed adverse pressure gradient for two Reynolds numbers (Re = 70 000, 150 000) and free-stream turbulence intensity levels (Tu = 1.5%, 2.5%). A wall-normal and two wall-parallel measuring planes located at different distance from the wall were considered. POD was applied to the entire PIV planes as well as on their sub-domains, showing the main flow features occurring in the different regions of the LSB. Then, the application of E-POD on different plane partitions revealed the existing correlation between the main dynamics observed in the forward part of the bubble and the breakup events occurring in the reattachment region. The E-POD modes computed in the breakup region resemble streaky structures when PIV snapshots are projected onto the POD eigenvectors of the near wall plane. Otherwise, Kelvin–Helmholtz rolls dominate the E-POD modes obtained by projection of the snapshot matrices on the basis computed in the plane located far from the wall. The main scales of the coherent structures highlighted by the E-POD modes were also characterized by means of the streamwise and spanwise autocorrelation functions of E-POD filtered fields. Data in this work clearly highlight the similarity properties of the main flow features observed in LSBs once scaled with the momentum thickness of the boundary layer at the separation position.

Comparative study on the aerodynamic performance of airfoil with boundary layer trip of various geometrical shapes

Performance of small scale wind turbine (SSWT) and miniature aerial vehicles (MAV) is always effected with Laminar separation bubble. The problem of a laminar separation bubble on the upper surface of an E216 airfoil operated at low Reynolds number (Re=100000) is investigated numerically. Incompressible steady two-dimensional simulation is carried out with Transition γ − Reθ turbulence model on the airfoil with a boundary layer trip (BLT). The performance of two different types of trips, namely, isosceles triangular (IT) and right-angled triangular (RA) is compared with that of the airfoil with a rectangular (RT) trip. The trip locations used are, 17% of the chord for location-1 and 10% of the chord for location-2 from the leading edge of the airfoil, while the trip heights selected are 0.3 mm, 0.5 mm, 0.7 mm, and 1 mm. Results showed that the boundary layer trip significantly affected the laminar separation and improved the aerodynamic performance of the airfoil. Maximum improvement in drag by 17.41% and corresponding lift to drag ratio by 10.86% are obtained for the isosceles trip at location-2 for the angle of attack of 6°. There is no observable advantage for isosceles and right-angled triangular trips over rectangular trips. Considering the geometrical complexity in fabrication, the rectangular trip is recommended.

A bionic noise reduction strategy on the trailing edge of NACA0018 based on the central composite design method

New innovative approaches to reduce noise are of great significance in engineering. Taking the typical NACA0018 airfoil as the study target, the serrated treatment of the trailing edge is carried out with the bionic noise reduction technology. To boost the design efficiency and clarify the distributing laws between the design parameters and the noise performances, a special investigation about optimizing the NACA0018 airfoil’s serrated trailing edge is implemented here. Firstly, a united parametric approach to represent various design schemes is proposed. Then, with the integration of the Computational Fluid Dynamics/Computational Acoustics analysis, Central Composite Design, and the Respond Surface Method, a bionic noise reduction strategy is established. The findings can be gotten as: taking into account the existence of the casing or wall near the span-end of the airfoil, the newly defined install location adjustment factor has an influence on the reduction noise, and the install location adjustment factor mainly affects the near-wall flow field of the serrated trailing edge; the optimal design scheme is obtained successfully, which can reduce the overall sound pressure level by about 2 dB in relative to the target; through the mechanism analysis of noise reduction, it can be found that the serrated trailing edge can suppress the development of the laminar separation bubbles, so the noise is reduced.

Laminar separation bubble dynamics and its effects on thin airfoil performance during pitching-up motion

This article reports an investigation into dynamic characteristics of the laminar separation bubbles (LSBs) associated with aerodynamic loads unsteadiness of a cambered thin airfoil in pitching-up motions at low Reynolds number flows. Unsteady Reynolds-averaged Navier–Stokes (URANS) simulations were conducted for a 4%c cambered thin airfoil at Reynolds number of 30,000 and 60,000. The airfoil pitches up from 0° to 25°angles of attack at dimensionless pitch rate [Formula: see text] of 0.0398 and 0.0199. The [Formula: see text] SST [Formula: see text] turbulence transition model was used to account for the effect of transition on LSBs’ development. The LSBs are shown to evolve in their shape and size during the pitching motion. The influence of the LSBs on the airfoil upper surface during pitching motion continues to a higher incidence in comparison with that under static conditions before developing into a fully detached flow. Vortex merging is observed in the rear part of the LSBs in the turbulent portion for a Reynolds number of 30,000. At Reynolds number 60,000, the changing of the LSB length during pitching-up motion is similar to that of steady cases, except a delayed transition is observed as incidence increases. The results show further insight into the dynamic characteristics of the LSBs and their relation to the aerodynamic performance of the airfoil.

Numerical Study on the Aerodynamic Characteristics of the NACA 0018 Airfoil at Low Reynolds Number for Darrieus Wind Turbines Using the Transition SST Model

A symmetrical NACA 0018 airfoil is often used in such applications as small-to-medium scale vertical-axis wind turbines and aerial vehicles. A review of the literature indicates a large gap in experimental studies of this airfoil at low and moderate Reynolds numbers in the previous century. This gap has limited the potential development of classical turbulence models, which in this range of Reynolds numbers predict the lift coefficients with insufficiently accurate results in comparison to contemporary experimental studies. Therefore, this paper validates the aerodynamic performance of the NACA 0018 airfoil and the characteristics of the laminar separation bubble formed on its suction side using the standard uncalibrated four-equation Transition SST turbulence model and the unsteady Reynolds-averaged Navier-Stokes (URANS) equations. A numerical study was conducted for the chord Reynolds number of 160,000, angles of attack between 0 and 11 degrees, as well as for the free-stream turbulence intensity of 0.05%. The calculated lift and drag coefficients, aerodynamic derivatives, as well as the location and length of the laminar bubble quite well agree with the results of experimental measurements taken from the literature for validation. A sensitivity study of the numerical model was performed in this paper to examine the effects of the time-step size, geometrical parameters and mesh distribution around the airfoil on the simulation results. The airfoil data sets obtained in this work using the Transition SST and the k-ω SST turbulence models were used in the improved double multiple streamtube (IDMS) to calculate aerodynamic blade loads of a vertical-axis wind turbine. The characteristics of the normal component of the aerodynamic blade load obtained by the Transition SST approach are much better suited to the experimental data compared to the k-ω SST turbulence model.

Numerical Investigation of Tonal Trailing-Edge Noise Radiated by Low Reynolds Number Airfoils

A high-fidelity computational analysis carefully validated against concurrently obtained experimental results is employed to examine self-noise radiation of airfoils at transitional flow regimes, with a focus on elucidating the connection between the unsteady behavior of the laminar separation bubble (LSB) and the acoustic feedback-loop (AFL) resonant interactions observed in the airfoil boundary layers. The employed parametric study examines AFL sensitivity to the changes in the upstream flow conditions and the airfoil loading. Implicit Large-Eddy Simulations are performed for a NACA-0012 airfoil in selected transitional-flow regimes for which experimental measurements recorded characteristic multiple-tone acoustic spectra with a dual ladder-type frequency structure. The switch between the tone-producing and no-tone-producing regimes is traced to the LSB size and position as a function of the flow Reynolds number and the airfoil angle of attack, and further substantiated by the linear stability analysis. The results indicate a strong multi-tonal airfoil noise radiation associated with the AFL and attributed to the switch from the slowly-growing Tollmien–Schlichting to the fast-growing Kelvin–Helmholtz instabilities occurring in thin LSB regions when those are localized near the trailing-edge (TE) on either side of the airfoil. Such a process eventually results in the nonlinearly saturated flapping vortical modes (“rollers”) that scatter into acoustic waves at the TE.

High-resolution direct numerical simulations of flow structure and aerodynamic performance of wind turbine airfoil at wide range of Reynolds numbers

The objective of this study is to develop direct numerical simulations (DNS) to investigate the aerodynamic performance, transition to turbulence, and to capture the laminar separation bubble occurring on a wind turbine blade. Simulations are conducted with spectral/hp element method to investigate the details of flow separation bubble over wind turbine blades with NACA-4412 airfoil at wide range of design parameters. This airfoil is chosen because recent studies have shown that it is challenging to capture the details of the flow instabilities and pressure fluctuations in the separated shear layer of wind turbines by experimental methods. Furthermore, owing to more accurate development of DNS, the separated bubbles at high Reynolds numbers are captured. The results show that the vortex structures shed from the trailing edge of the airfoil by raising the angle of attack (α). Consequently, the fully turbulent flow develops downstream of the trailing edge (Karman vortex). Moreover, the pressure fluctuation significantly increased by raising α. However, some rolling up of the flow structures, similar to Kelvin–Helmholtz rolls, on the pressure surface near the trailing edge, are observed at α>12°. The separation point was delayed from Xsep/C = 0.19 to 0.58 by decreasing α from 16 to 0 at Re = 5 × 104.

Simulation and characterization of the laminar separation bubble over a NACA-0012 airfoil as a function of angle of attack

Airfoils operating at low Reynolds number have a proclivity to induce a laminar separation bubble (LSB) on their upper surface. We investigate the effects of angle of attack on the characteristics of the LSB and the flow field around a NACA0012 airfoil. Large-eddy simulations show that there are three distinct angle-of-attack regimes: a pre-stall attached flow, a near stall separating-attaching flow, and a full stall separated flow without reattachment.

Effect of Ice accretion on the aerodynamic characteristics of wind turbine blades

Cold regions with high air density and wind speed attract wind energy producers across the globe exhibiting its potential for wind exploitation. However, exposure of wind turbine blades to such cold conditions bring about devastating impacts like aerodynamic degradation, production loss and blade failures etc. A series of wind tunnel tests were performed to investigate the effect of icing on the aerodynamic properties of wind turbine blades. A baseline clean wing configuration along with four different ice accretion geometries were considered in this study. Aerodynamic force coefficients were obtained from the surface pressure measurements made over the test model using MPS4264 Simultaneous pressure scanner. 3D printed Ice templates featuring different ice geometries based on Icing Research Tunnel data is utilized. Aerodynamic characteristics of both the clean wing configuration and Ice accreted geometries were analysed over a wide range of angles of attack (α) ranging from 0° to 24° with an increment of 3° for three different Reynolds number in the order of 105. Results show a decrease in aerodynamic characteristics of the iced aerofoil when compared against the baseline clean wing configuration. The key flow field features such as point of separation, reattachment and formation of Laminar Separation Bubble (LSB) for different icing geometries and its influence on the aerodynamic characteristics are addressed. Additionally, attempts were made to understand the influence of Reynolds number on the iced-aerofoil aerodynamics.