Laminar Separation Research Articles

Study of the influence of turbulent flows on the aerodynamic performance of fixed-wing aircraft type UAV MALE

A study of the air behavior in the near-wake region of the NACA0012 airfoil infinite wing model at low Reynolds numbers was conducted. The study consisted of comparing the results obtained from experiments with those from numerical simulations. The numerical results of the velocity profiles in the near-wake region showed good agreement with those provided by the PIV system for most of the simulations and those in the literature. In addition, boundary layer separation and reattachment forming a laminar separation bubble along the airfoil was captured using simulations in Ansys Fluent and compared with a literature database to better confirm the simulation results. The UAVs have a complex cross-section shape of the fuselage. The second part of this paper explores the aerodynamic consequences of its shape, at the operational flight ceiling, in a comparative approach for three different fuselage designs.

Features of laminar separation bubble subjected to varying adverse pressure gradients

This article describes the spatial development of a laminar separation bubble (LSB), its transition, and eventual breakdown under the influence of adverse pressure gradients (APGs) similar to those experienced by low-pressure turbine blades. The investigation combines a comprehensive experimental approach with a well-resolved large eddy simulation (LES). The streamwise pressure gradients were varied by manipulating the upper wall within the test section. The Reynolds number (Re), based on the plate length and inlet velocity, was 0.2 × 106 with a freestream turbulence intensity of 1.02%. The particle image velocimetry (PIV) and hotwire data were used to illustrate the vortex dynamics, growth of perturbations, and intermittency. The onset and end of transition progressively shift upstream, resulting in a reduction of the laminar shear layer length and bubble length with increasing APG. Interestingly, the flow features exhibit self-similarity in velocity profiles and the growth rate of velocity fluctuations when normalized against the bubble length. The formation of two-dimensional Kelvin–Helmholtz (K–H) rolls is apparent in the beginning, resulting in the selective amplification of frequency and exponential growth of fluctuations. Linear stability theory explains the most amplified frequency and phase speed of convective vortices, apart from the growth of disturbances. Analysis of LES data reveals intricate inviscid–viscous interactions that trigger shear layer breakdown. In brief, evolving perturbations within the braid region of vortices in the latter half interact with the advecting K–H rolls, culminating in the breakdown and the onset of turbulent flow downstream.

Flow Confinement Effects on sUAS Rotor Noise

The goal of this paper is to investigate the effects of flow confinement on the noise generated by a small unmanned aerial system rotor measured and simulated in a partially closed test room. More specifically, the influence of the flow recirculation in the test room and the consequent blade turbulence interaction on both the tonal unsteady loading noise and the broadband laminar separation noise components are scrutinized. The study also aims at improving our prediction capabilities of drone rotor noise, and deepening our fundamental knowledge of rotor aeroacoustics in transitional boundary layer regimes. Numerical simulations are conducted by means of a scale-resolving wall-modelled lattice Boltzmann solver to calculate the unsteady flow and the noise generated by the same reference configuration used in our previous works, which consists in a rotor operated at a chord-based Reynolds number of about 70000 at 75% of the tip radius, both in hover and axial flow conditions. Results for a rotor in pristine flow conditions are compared to results obtained by simulating the rotor in a confined environment that reproduces the A-Tunnel test chamber of Delft University of Technology. One one hand, our study reveals that previously observed discrepancies between measured and predicted noise spectra in hover conditions are not due to the influence of the flow confinement on the laminar separation mechanisms, as hypothesized in our previous study, rather to a lack of numerical resolution. On the other hand, the new results reveal that, in hover conditions, the dominant effect of the flow recirculation is to increase the unsteady loading tonal noise components, whereas, in axial flow conditions, no significant effect due to the test environment is observed. Some of our conclusive observations are supported by results obtained using a source identification technique herein presented for the first time.

Numerical Study on Aerodynamic and Aeroacoustic Performances of Bioinspired Wings

Noise control has become one of the key issues to be considered in modern aeronautical machinery design. Many efforts have been devoted to noise reduction of airfoils and wings, including traditional flow control methods. In fact, some animals in wild nature exhibit superior aerodynamic and aeroacoustic performance, providing novel ideas for solving this engineering problem. In this research, bionic technology is used to obtain quiet and efficient wing. Inspired by the owl’s wing, we propose two bionic configurations, one coupled with leading edge waves and trailing edge serrations. The Large Eddy Simulation and the Ffowcs-Williams and Hawkings equation is applied to simulate the aerodynamic and aeroacoustic characteristics of wings at low-Reynolds number flow. Numerical results demonstrate that the bioinspired wings have excellent aerodynamic performances and remarkable lower overall sound pressure level compared to NACA 0016 which has similar relative thickness. In addition, the unsmooth structure of leading edge waves and trailing edge serrations provide an additional 4.27 dB noise suppression effect, with little impact on aerodynamic characteristics at small angle of attack. The detailed analysis reveals that, due to the special owl-based profile, the flow around two bioinspired wings is mainly turbulent on the upper and lower surfaces, and no laminar separation bubble is detected at the trailing edge. Moreover, the unsmooth structure modifications successfully weaken the scale and scope of coherent vortex structures. These factors contribute to reducing the associated pressure fluctuation, thereby controlling the aeroacoustic noise of wing. Consequently, a coupled bionic wing is presented with the excellent aerodynamic and aeroacoustic characteristics. The conclusions are envisioned to be beneficial to the design of new generation low-noise aeronautical machinery.

Effects of number of blades on propeller noise

This paper presents the effect of the number of blades on the aeroacoustic characteristics and aerodynamic performance of propellers utilized in urban air mobility vehicles through an extensive experimental study. A detailed parametric study is conducted for two, three, and five-bladed propellers. Investigations are carried out under static thrust conditions and incoming axial flow velocities for various rotational speeds. The spectra and directivity of the far-field noise are studied in detail for broadband and tonal components. Results show that two-bladed propellers produce higher tonal noise levels than other blade numbers for the same tip speed, while five-bladed propellers produce the highest broadband noise levels. The highest level of noise radiation is noted in the upstream and downstream directions for broadband noise, whereas tonal noise is radiated mainly in the plane of rotation. Results also indicate that high-frequency broadband humps are observed for higher advance ratios, likely caused by the formation of laminar separation bubbles, and that amplitude increases with blade numbers. The results of this study provide valuable insight for further investigations of propeller noise and aerodynamic performance.

Modeling Laminar Separation Bubbles by Field Inversion Method

This paper presents the application of a data-driven method for turbulence modeling. The main aim is to enhance the prediction capabilities of an existing Reynolds-averaged Navier–Stokes (RANS) model for laminar separation bubbles. The recently developed field inversion method has been chosen to infer corrections to the shear stress transport turbulence model. An innovative approach has been proposed to reduce the computational cost of the modeling procedure. The objective function makes use of the friction coefficient to compare data coming from the RANS model with reference large-eddy simulation data available in the literature. The performed test cases responded well to the optimization procedure which provided a correction field that boosts the turbulent kinetic energy within the bubble. The proposed procedure is the necessary building block to define a new turbulence model able to overcome the present limitations in simulating laminar separation bubbles.

Numerical study on the periodic control of supersonic compression corner flow using a nanosecond pulsed plasma actuator

This study investigates the effectiveness of a pulsed nanosecond dielectric barrier discharge (NSDBD) plasma actuator for flow control over a supersonic compression corner through numerical simulations. The effects of varying applied voltages, repetitive frequencies, and activated locations of the plasma actuator are examined under large-scale laminar flow separation conditions around a compression corner. The unit Reynolds number and Mach number are 7.8 × 106 m−1 and 4, respectively. The results indicate that the discharge induces a pressure rise and leads to misalignment between the pressure gradient and density gradient in the residual heat region. Because of the interaction between the supersonic freestream and the actuation-induced shock/compression flow, convection, compressibility of the fluid element, and baroclinicity of the residual heat region collectively lead to the formation of an induced spanwise vortex, which in turn enables momentum migration. The induced vortex disrupts the initial flow structures and entrains high-energy fluid from the main flow into the boundary layer, promoting momentum mixing between the main flow and the separated flow, which increases the energy of the boundary layer to resist the adverse pressure gradient. The time-averaged flow structures imply that it is possible to totally eliminate the flow separation near the supersonic compression corner. For aerodynamics on the surface, the normal force produces a pitching moment that can be potentially utilized to control the body's orientation and trajectory. Additionally, total drag on the surface can be reduced by 5 %. This suggests that choosing the most appropriate position based on its local fluid characteristics can strongly increase the control effectiveness.

Numerical study of an acoustic field effect on boundary layer transition mechanism

The laminar-to-turbulent transition mechanism is fundamentally important for the concepts aiming at reducing CO2 emissions and noise pollution. This research is focused on understanding the acoustic excitation effect on the transition mechanism with a laminar separation bubble. The transition over the suction side of the NACA 0012 airfoil at an angle of attack of 2 degrees and Reynolds number of 2 × 105 with zero free-stream turbulence intensity is considered. The investigation is performed based on three flow cases obtained for a smooth and wavy surface modified airfoils using the large-eddy simulation method. The acoustic excitation is controlled by keeping or removing a naturally occurring acoustic feedback loop by introducing surface waviness on the airfoil’s pressure side. The flow solution for the standard NACA 0012 airfoil provides the acoustic excitation case, whereas the airfoil with the wavy surface on the pressure side provides the case without acoustic excitation. The results confirm that a shorter separation bubble and earlier laminar-to-turbulent transition can be triggered in response to the acoustic excitation. Furthermore, the coupling of the acoustic field and hydrodynamic instabilities is investigated. Finally, it is shown by referring to experimental data that the Kelvin-Helmholtz instabilities inside the separation bubble are amplified by the acoustic excitation.

Numerical Study of the Influence of the Critical Reynolds Number on the Aerodynamic Characteristics of the Wing Airfoil

The paper reports the results of a study concerned with the influence of the size of the leading edge laminar bubble on the aerodynamic characteristics of the HGR01 airfoil. The completely turbulent and transient flows are considered. The mechanism of the appearance and interaction of laminar and turbulent flow separation near the leading and trailing edges of the airfoil is studied in detail. In the paper, the dependence of aerodynamic forces on the critical Reynolds number for the HGR01 airfoil is discussed. It has been established that the separation bubble at the leading edge can only be obtained using the laminar–turbulent transition model. Fully turbulent models are not able to show this feature of the airfoil flow. Graphs of the lift coefficient as a function of the critical Reynolds number, as well as the pressure distribution as a function of the size of the laminar bubble, are shown.

Dual transition scheme on [formula omitted]-equation model

The proposed model integrates two transition mechanisms into the turbulent kinetic energy (k-equation) turbulence framework, assisted by transition representatives. These include a “flow-structure-adaptive” stress-intensity parameter, which induces the pre-transitional/pseudo-laminar state before transition, and an intermittency factor, which facilitates the prediction of flow transition onset and completion with a feasible growth rate and a logical transition length. The algebraic transition model introduces new functions and correlations, which are based on theoretical and experimental evidence. These elements stimulate multiple transition phenomena in a suitable and credible way due to their reliance on local flow information for initiating and controlling the transition growth rate. With the employment of an algebraic closure for the dissipation rate, the k-equation directly anticipates the free-stream turbulence intensity instead of the free-stream “eddy-to-laminar” viscosity ratio (Rμ∞). This approach avoids the “trial-and-error” inconsistency typically associated with most correlation-based and physics-based transition models when initiating appropriate computations. The independence of the algebraic transition model from Rμ∞ offers substantial benefits in predictive capability over traditional transition models. However, the quality of performance might fluctuate, contingent on the closure approximations adopted. This detail is of immense importance for accurately depicting the pertinent physical characteristics of the flow, such as bypass, natural, and separation-induced transitions. Numerical results indicate that the current model, whether equipped with algebraic transition additives or not, aligns well with both existing experimental data and commonly used transition and non-transition models. The new transition model has decent agreement with the transitional boundary layer on a flat plate, and transitional flow over an airfoil with laminar and turbulent separation bubbles.

Investigation on transition characteristics of laminar separation bubble on a hydrofoil

The phenomenon of water–jet pump stall can be ascribed to the development of blade boundary layer separation with the transition process playing a significant role in this separation. The hydrofoil is usually used as a simplified model of the water–jet pump impeller blade, and its flow field characteristics have important reference values for analyzing the impeller flow. Based on the transition model and the dynamic mode decomposition method, this article presents the results of a study that was carried out on the stall characteristics of the NACA0009 blunt trailing edge hydrofoil. The transition characteristics of hydrofoil surfaces at different angles of attack (AoA)and Chord-based Reynolds numbers (ReL) are obtained. The hydrofoil boundary layer transition is dominated by natural transition as the AoA is less than 4°, while the transition is dominated by leading-edge separation-induced transition as the AoA is greater than 4°. The investigation yields the dynamic properties of the LSB (Laminar Separation Bubble) as the AoA is varied. The phenomenon known as the deep stall is distinguished by the movement of the stall vortex toward the upstream direction near the trailing-edge region, where it merges with the LSB in the leading-edge region. This phenomenon leads to oscillations in the lift and drag coefficients. The relationship between the LSB and the trailing-edge stall vortex is established using DMD (Dynamic Mode Decomposition) methods. As the phenomenon of the deep stall occurs, it can be observed that the modal energy of the leading-edge LSB is comparatively higher than the modal energy of the trailing-edge stall vortex, inducing the dominant role of the LSB and the movement toward the trailing-edge region and, consequently, the phenomenon of trailing-edge vortex shedding in the hydrofoil. The findings of this study could be guidance for the design of fluid machinery blades.

Airfoil optimization for rotors operating in the ultra-low Reynolds number regime

Ultra-low Reynolds number aerodynamics has gained significant attention in recent times, primarily due to the proliferation of micro-aerial vehicles and their utilization in low-density environments such as the Martian atmosphere and high altitudes in Earth's atmosphere. The Martian atmosphere presents a unique combination of low Reynolds numbers and high subsonic Mach numbers, necessitating the use of unconventional airfoil designs in this regime. This study focuses on optimizing airfoils for rotors operating in the compressible ultra-low Reynolds number regime. We demonstrate the capability of XFOIL, a panel method code incorporating an integral boundary layer formulation, to accurately predict airfoil loads when the flow remains attached. We determine the optimal camber and maximum camber positions by analyzing the four digits, two percent thickness, National Advisory Committee for Aeronautics XX02 airfoil family using XFOIL. Subsequently, we employ a multi-objective optimization approach, utilizing a Class Shape Transformation airfoil parameterization to maximize lift and minimize drag. We select several points from the resulting Pareto front and evaluate their performance through unsteady compressible Navier–Stokes simulations. Our findings reveal that incorporating sharp leading-edge variations in these airfoil designs enhances the peak efficiency by over 10%, primarily attributable to the development of laminar separation bubbles on the suction side of the airfoils. Importantly, these modified airfoils maintain favorable performance at low angles of attack.

Large eddy simulations and experiments on low-Reynolds-number laminar separation control for low-pressure turbine blades with oblong-dimpled surface

Modern high-performance low-pressure turbines (LPT) usually encounter flow problems at low Reynolds numbers, especially for high-altitude UAVs under cruise conditions. Flow separation may occur on the blade suction side, seriously degrading the turbine's aerodynamic efficiency. The present study investigates a passive laminar flow separation control method based on the oblong dimple surface structures for their more potent flow acceleration abilities. Linear cascade tests and large eddy simulations are conducted for a typical LPT blade profile, T106A, at the Reynolds numbers of 25,000, 50,000, and 100,000 based on the inlet velocity and axial chord length. Results show that applying the oblong dimples can apparently reduce the reverse flow region and improve the aerodynamic performance at low Reynolds numbers. The experimental results achieve a maximum total pressure loss reduction of 34.7% at a low Reynolds number of 25,000. The numerical simulations show the local flow acceleration after the oblong dimples and demonstrate two working modes. At low Reynolds numbers, the oblong dimples accelerate the fluid above, generating local high-speed flows, and the uneven velocity distributions in the spanwise direction feed the deformation and fragmentation of the downstream separating vortex. While at high Reynolds numbers, the oblong dimples act as vortex generators, leading to flow transition on the blade suction surface.

Flow structure and vortex dynamics of blade surface transitional flow in the near-tip region of a compressor cascade

The mean flow topology and vortex dynamics of the blade surface transitional flow in the near-tip region of a C4 compressor blade were investigated using particle image velocimetry measurements with two configurations. The experiment was conducted in a compressor cascade at a chord Reynolds number of 24 000 and an incidence angle of 0°, and a laminar separation bubble was detected on the aft portion of the blade. In the half-span region of the blade, the separation is essentially two-dimensional without reattachment. The vortex dynamics are dominated by the periodic shedding of separated shear layer vortices and their interaction with the trailing edge vortices. The progressive spanwise evolution in the flow structures and vortex dynamics occur near the blade tip (70%–80% blade height), leading to an advanced, thinner separation. In contrast, the tip leakage vortex dominated region is restricted to approximately 20% of blade height from the blade tip. In this region, secondary flow effects are intensive enough to prevent laminar separation. Between the above two regions, there is a transition region (90% blade height), where the influence of the tip flow on the blade surface flow is relatively slight that merely suppresses the vortex shedding of the separated boundary layer, nor the whole shear layer. In the transition region, the velocity fluctuations are significantly reduced.

Effect of wavy leading edges on airfoil tonal noise

This work presents numerical studies on the effect of wavy leading edges on airfoil instability tonal noise. Large eddy simulations are conducted for NACA0012 (National Advisory Committee for Aeronautics) airfoils with straight and wavy leading edges. The far-field aerodynamic noise is predicted using the Ffowcs Williams and Hawkings acoustic analogy theory. The inflow Mach number is approximately 0.17 with an angle of attack of 4.3°, and the chord based Reynolds number is 600 000. The present numerical method is first validated by existing numerical results and a semi-empirical model for the straight baseline airfoil. Then, the effects of wavy leading edges on the aerodynamic performance, aeroacoustic performance, and noise reduction mechanisms are discussed in detail. The wavy leading edge is found to be detrimental to the mean aerodynamic performance with a decreased lift and increased drag. However, the lift fluctuations are significantly reduced. The instability tonal noise and its harmonic are totally removed by the wavy leading edges, while an extra broadband hump appears at the middle frequency. The low frequency and high frequency noise are also increased by the wavy airfoil. The laminar separation bubble is removed, and the flow separation is decreased by the wavy airfoil in the vicinity of the trailing edge. The pressure fluctuations around the trailing-edge are considerably diminished, and the spanwise coherence spectra are also significantly reduced by the wavy airfoil. Furthermore, it is noteworthy that the coherence reduction spectrum at the source is almost consistent with the far-field noise reduction spectrum.

Aerodynamic Performance of an Unmanned Aerial Vehicle Wing For Varied Wing Geometric Parameters

The aerodynamic performance characteristics of a small unmanned aerial vehicle wing are evaluated computationally at a moderate Reynolds number of Re = 2.8 x 105. The flow exhibits laminar-turbulent transition and features a laminar separation bubble that exists for a substantial portion of the wing span. The flow near the wing tip is three-dimensional and exhibits complex flow patterns due to the interaction of the wing tip vortex with the laminar separation bubble. A series of systematic parametric studies are conducted by varying the wing geometric parameters like the span of the rectangular inboard wing, sweep of the outboard wing and wing dihedral angle. From the parametric studies, it is found that the wing with an outboard sweep of 17° has a 4.54 % improvement in endurance factor compared to the baseline wing at an angle of attack of 5°. The drag of this wing is 19.2 drag counts lower compared to the baseline wing. Further, flow transition is delayed for this wing. The present results show that wing sweep has a significant influence on the improvement of aerodynamic performance of unmanned aerial vehicle wings at moderate Reynolds numbers.

Capturing the drag crisis in the flow around a smooth cylinder using a hybrid RANS-LES model on coarse meshes

A hybrid strategy combining Reynolds-averaged Navier–Stokes (RANS) and large eddy simulation (LES) methods is nowadays seen as an efficient way to simulate turbulent flows of practical relevance. In this work, the scale-resolving hybrid (SRH) model proposed by Manceau (2018) is compared with a conventional unsteady RANS model, and LES and experimental data from the literature for the flow around a smooth cylinder in the flow regime around the drag crisis. Based on a temporal filtering formalism, this approach has seen limited testing for turbulent separated flows. The drag crisis phenomenon is dominated by complex near-wall physics and is challenging to simulate. The predictive accuracy and the robustness to mesh coarsening for the SRH model are assessed for this test case, with the aim to demonstrate that this hybrid approach can be a credible cost-saving alternative to LES for separated turbulent flows. The meshes considered in this numerical study are far coarser than the ones used in the LES reference data, yet, the results for the time mean drag are found in good agreement with the reference data. Other features of the flow, such as the presence and sizes of the laminar separation bubbles and the consequent magnitude of the time mean lift are not as well captured. In general, the qualitative behaviour of the SRH model is good when considered in the context of questions previously raised in the literature about hybrid models. The mesh savings are achieved by coarsening the spatial resolution in the wake, whereas the resolution required in the near-wall area and shear layers remains high, though much reduced compared to the reference LES data. The key point taken from this study is that the SRH model is an attractive option to produce higher-fidelity data on coarse meshes.

Excitation of a Separated Flow by a Series of Protuberances Near the Leading Edge of a Model Aerofoil

Abstract This article documents the manifestation of a shear layer under the excitation of a series of hemispherical protuberances near the leading edge of a constant-thickness airfoil. The experiments are performed at a Reynolds number of 1.6 × 105 based on the chord length and inlet velocity, where freestream turbulence is 1.2%. The hotwire and particle image velocimetry data are analyzed to appreciate the flow feature, illustrating the growth of perturbations, vortex dynamics, intermittency, and spectral response. A laminar separation bubble (LSB) appears near the leading edge for a smooth surface, and the shear layer is inviscidly unstable. The evolution of the shear layer significantly changes with a series of protuberances. The breakdown of the shear layer occurs almost immediately, triggering local turbulence resulting in a considerable reduction of the bubble length. However, a separation bubble of varying spanwise lengths is formed in this case. Although the power spectra of velocity fluctuations reveal the selective amplification of frequencies even with protuberances, the immediate augmentation of turbulence followed by faster decay suggests the transient growth of turbulence. The study has documented insight into features of a separation bubble subjected to leading-edge perturbations and might influence future studies on separation control over an airfoil.

Instability and transition onset downstream of a laminar separation bubble at Mach 6

Instability measurements of an axisymmetric, laminar separation bubble were made over a sharp cone-cylinder-flare with a$12^{\circ }$flare angle under hypersonic quiet flow. Two distinct instabilities were identified: Mack's second mode (which peaked between 190 and 290 kHz) and the shear-layer instability in the same frequency band as Mack's first mode (observed between 50 and 150 kHz). Both instabilities were measured with surface pressure sensors and were captured with high-speed schlieren. Linear stability analysis results agreed well with these measured instabilities in terms of both peak frequencies and amplification rates. Lower-frequency fluctuations were also noted in the schlieren data. Bicoherence analysis revealed nonlinear phase-locking between the shear-layer and second-mode instabilities. For the first time in axisymmetric, low-disturbance flow, naturally generated intermittent turbulent spots were observed in the reattached boundary layer. These spots appeared to evolve from shear-layer-instability wave packets convecting downstream. This work presents novel experimental evidence of the hypersonic shear-layer instability contributing directly to transition onset for an axisymmetric model.

Piezoelectric Based Adaptive Wing for Flow Control

An attempt is made to develop a piezoelectric based adaptive wing concept that can be used in Micro Air Vehicle (MAV) and Unmanned Air Vehicle (UAV) applications. Laminar boundary layer separation is considered to be a significant problem in low Reynolds number vehicles like MAV. Recently, periodic excitation has gained much interest in the control of flow separation through energizing the boundary layer. In the present study, a piezoelectric bimorph actuator has been used to develop a mechanism that will deflect or vibrate a segment of the wing to produce the required thickness variation and achieve flow control. A Finite Element model of the adaptive wing is developed in ANSYS and numerical results are compared with experiments. A five percent increase in the camber has resulted in 4.39 percent additional lift at higher angles of attack.