Laminar-turbulent Transition In Boundary Layer Research Articles

Description of laminar-turbulent transition of an airfoil boundary layer measured by differential image thermography using directed percolation theory

Description of laminar-turbulent transition of an airfoil boundary layer measured by differential image thermography using directed percolation theory

Numerical simulation of laminar-turbulent transition in a supersonic boundary layer under the action of acoustic disturbances

In low disturbance environment, laminar-turbulent transition (LTT) is associated with excitation of unstable normal modes of small initial amplitudes (receptivity problem). These modes grow exponentially to a critical amplitude in accord with the linear stability theory (LST) and trigger the nonlinear breakdown. In the current study numerical simulation of laminar-turbulent transition (LTT) in the boundary layer on the upper surface of a flat plate in a supersonic free stream of Mach number M∞=3 and Reynolds number Re∞=2 × 107 is carried out. The flow is perturbed by fast or slow acoustic waves of low intensity, which excite unstable waves of the first mode. The latter have frequencies corresponding to the integral amplification N ≈ 9.16 typical for flight conditions. Two cases are considered: the angle of attack AoA=0° at which acoustic waves pass through a weak shock induced by viscous-inviscid interaction; AoA=5° at which acoustic waves pass through the expansion fan emanating from the plate leading edge. A holistic modeling of all stages of the transition from receptivity to the birth of turbulent spots is performed using direct numerical simulations. Linear stability theory is used to interpret the results. Feasibility of practical implementation of the amplitude method for predictions of the LTT onset in the considered and similar cases is discussed. It has been shown that to realize the amplitude method in the considered and similar cases, it is sufficient: to calculate the initial amplitudes of instability in a small vicinity of the leading edge and the propagation of instability from the leading edge to the station where the instability reaches the critical amplitude u′max≈4%. Analysis of known numerical and experimental data showed that this criterion weakly depends on the local Mach number in the range 0<Me<7, and it is acceptable for practical predictions of the transition onset points.

Miniaturization and Model-Integration of the Optical Measurement System for Temperature-Sensitive Paint Investigations.

The temperature-sensitive paint (TSP) method, an optical measurement technique, is used for qualitative skin friction visualizations in a wide variety of aerodynamic applications. One such application is the visualization of the laminar-turbulent boundary-layer transition. Optical access to the surface of interest is mandatory for the measurement system, which consists of scientific cameras and LEDs. But the optical access to the area of interest is often impeded by the available windows of the wind tunnel and the wind tunnel model itself, reducing the field of view and the spatial resolution. In some cases, it is of interest to increase the flexibility of the installation of the optical measurement system by reducing its physical dimensions and placing the installation inside the plenum. The DLR Swept flat PlatE Cross-flow TRAnsition (SPECTRA-A) configuration was selected to investigate the influence of two-dimensional steps on the cross-flow-induced boundary layer transition by means of TSP, as part of the EU project Clean Sky 2. The SPECTRA-A configuration consists of two main elements: a flat plate and a displacement body mounted within a very close distance of each other, creating a narrow gap between the two elements. The surface of interest is the area on the flat plate facing the displacement body. The narrow gap limits the utilization of an external camera setup due to poor optical access. A new optical setup consisting of four miniature CMOS machine-vision cameras and five miniature high-power LEDs was integrated into the displacement body. The characteristics of the camera system were analyzed in laboratory tests, establishing that the miniature CMOS machine-vision cameras are suitable for qualitative TSP skin friction visualizations. This was confirmed by successfully measuring the laminar-turbulent boundary-layer transition on the SPECTRA-A configuration. The integrated TSP system is capable of resolving even small variations of the transition location caused by changing the amplitude of the stationary cross-flow instability. The quality of the TSP visualization with the integrated optical system allows for the measurement of the transition location and the wavelength of the stationary cross-flow instability. Overall, a cost-effective TSP visualization system with small space requirements was developed and tested for future applications in wind tunnel models, model support, or side walls of wind tunnels.

Effect of Small Angles of Attack on Turbulence Generation in Supersonic Boundary Layers on Swept Wings

We present the new (for Mach numbers М = 3 and 3.5) and generalizing (for Mach numbers from 2 to 4) results of experimental investigations on the effect of small angles of attack on laminar-turbulent transition in the supersonic boundary layer on a swept wing with the leading-edge slip angle of 72°. The angle-of-attack variation has a strong effect on the transition Reynolds number. The transition Reynolds number decreases with increase in the Mach number. The measurements were carried out by means of a constant-temperature hot-wire anemometer using the proven procedure of determining the transition location. The eN method is used for the first time for numerically estimating the transition Reynolds numbers in the supersonic boundary layer on a swept wing with the leading-edge slip angle of 72°. The growth of the amplitudes of the steady and unsteady modes of the boundary layer crossflow are calculated in accordance with the linear stability theory, within the framework of the Lees–Lin system of equations. The numerical results indicate that, in accordance with the experimental results, laminar-turbulent transition in the boundary layer on the model swept wing is governed by the growth of stationary modes of the crossflow instability.

Effect of Small Angles of Attack on Turbulence Generation in Supersonic Boundary Layers on Swept Wings

We present the new (for Mach numbers М = 3 and 3.5) and generalizing (for Mach numbers from 2 to 4) results of experimental investigations on the effect of small angles of attack on laminar-turbulent transition in the supersonic boundary layer on a swept wing with the leading-edge slip angle of 72°. The angle-of-attack variation has a strong effect on the transition Reynolds number. The transition Reynolds number decreases with increase in the Mach number. The measurements were carried out by means of a constant-temperature hot-wire anemometer using the proven procedure of determining the transition location. The eN method is used for the first time for numerically estimating the transition Reynolds numbers in the supersonic boundary layer on a swept wing with the leading-edge slip angle of 72°. The growth of the amplitudes of the steady and unsteady modes of the boundary layer crossflow are calculated in accordance with the linear stability theory, within the framework of the Lees–Lin system of equations. The numerical results indicate that, in accordance with the experimental results, laminar-turbulent transition in the boundary layer on the model swept wing is governed by the growth of stationary modes of the crossflow instability.

On the laminar–turbulent transition mechanism on megawatt wind turbine blades operating in atmospheric flow

Abstract. Among a few field experiments on wind turbines for analyzing laminar–turbulent boundary layer transition, the results obtained from the DAN-AERO and aerodynamic glove projects provide significant findings. The effect of inflow turbulence on boundary layer transition and the possible transition mechanisms on wind turbine blades are discussed and compared to CFD (computational fluid dynamics) simulations of increasing fidelity (Reynolds-averaged Navier–Stokes, RANS; unsteady Reynolds-averaged Navier–Stokes, URANS; and large-eddy simulations, LESs). From the experiments, it is found that the transition scenario changes even over a single revolution with bypass transition taking place under the influence of enhanced upstream turbulence, for example, such as that from wakes, while natural transition is observed in other instances under relatively low inflow turbulence conditions. This change from bypass to natural transition takes place at azimuthal angles directly outside the influence of the wake indicating a quick boundary layer recovery. The importance of a suitable choice of the amplification factor to be used within the eN method of transition detection is evident from both the RANS and URANS simulations. The URANS simulations which simultaneously check for natural and bypass transition match very well with the experiment. The LES predictions with anisotropic inflow turbulence show the shear-sheltering effect and a good agreement between the power spectral density plots from the experiment and simulation is found in case of bypass transition. A condition to easily distinguish the region of transition to turbulence based on the Reynolds shear stress is also observed. Overall, useful insights into the flow phenomena are obtained and a remarkably consistent set of conclusions can be drawn.

Active transition control by synthetic jets in a hypersonic boundary layer

We investigate by direct numerical simulation the active control of laminar-turbulent transition in a hypersonic flat-plate boundary layer at a freestream Mach number of 5.86. The control mechanism is a synthetic jet. Based upon the linear stability theory of Mack, in hypersonic flow the important path to transition involves a high-frequency, second-mode fundamental resonance. Through systematic investigation, we reveal that the forcing the boundary layer with a synthetic jet at appropriate combinations of amplitude and frequency suppresses the second mode and delays transition. To gain physical insights into the major control mechanism, we employ the momentum potential theory (MPT) to analyze the flows with and without control. Essentially, the underlying control mechanism relies on an intriguing effect of the synthetic jet via generating the outward radiated wave structures, which are identified to split the upstream acoustic and vortical components. The splitting treatment presents the second-mode energy to drop sharply after the flow passes through the synthetic jet slot. The MPT source-term analysis reveals that the significantly suppressed near-wall source terms are responsible for suppressing the second mode downstream. Compared with the vortical and thermal source terms, the acoustic source term is found to be suppressed most. The kinetic budget analysis further reveals that the splitting treatment is related to the non-parallel effect and the nonlinear interaction.

Impact of a continuously extending wall on laminar-turbulent transition in subsonic boundary layers

Structural morphing is an efficient means to enable the design of aircrafts for diverse requirements, such as high endurance and high speed; however, there may be side effects on the aerodynamic performance, which needs to be considered before its implementation. In this paper, we particularly focus on the impact of a representative morphing structure, a continuously extending wall in the chordwise direction, on the laminar-turbulent transition in subsonic boundary layers. The large-Reynolds-number asymptotic approach is employed, and both the mean-flow distortion and the perturbation field are described by the triple-deck theory, which leads to a great reduction in the number of the controlling parameters and shows clearly their interaction mechanisms. Two relevant mechanisms, namely, the local receptivity and local scattering mechanisms, are considered, whose effects on the development of the boundary-layer instability modes, leading to a change in the transition onset eventually, are quantified systematically by solving numerically the high-dimensional linear equation system. The receptivity efficiency is greater near the lower-branch neutral frequency, while the scattering effect increases with the frequency monotonically. Both the receptivity and scattering calculations show good agreement with the linear predictions when the extending-wall speed is sufficiently low, but for a moderate extending speed, both the receptivity efficiency and the scattering efficiency increase superlinearly with the extending speed.

Scattering of Mack modes by solid-porous junctions in hypersonic boundary layers

Porous coating is regarded as an effective control strategy to delay the laminar-turbulent transition in hypersonic boundary layers, whose effects include not only the successive distortion of the Mack-mode growth rate but also the scattering effect appearing at the junctions between the solid and porous walls. The present paper focuses particularly on the latter mechanism by employing the harmonic linearized Navier–Stokes approach, which is confirmed to be sufficiently accurate as compared to the results obtained by direct numerical simulations. A transmission coefficient is introduced to quantitatively characterize the scattering effect, which is define as the ratio of the Mack-instability amplitude downstream of the junction to that upstream. For a solid-to-porous junction, the scattering effect leads to a stabilization of the majority of the second mode but destabilizes slightly the first mode. In contrast, the porous-to-solid junction plays the opposite role. The overall effect of a finite-length porous panel requires a combined consideration of the Mack-mode-growth-rate distortion and the scattering effect of the two junctions at both boundaries of the panel, which shows a critical frequency separating the destabilizing and stabilizing frequency bands. Additionally, a decrease in the wall temperature leads to an enhanced scattering effect on the Mack modes.

Influence of Roughness Trips on Near- and Far-Field Trailing-Edge Noise

Roughness trips are widely applied in wind-tunnel testing to hasten the laminar–turbulent boundary-layer transition. It is well known that they influence the boundary-layer transition and its development. However, their influence on aeroacoustic quantities (in particular, on the airfoil trailing-edge noise) is not yet fully understood. This paper discusses the relationship between the trip height and the boundary-layer development and radiated trailing-edge noise. Wind-tunnel experiments were performed with zigzag strips and grit elements placed on the surface of a NACA 0012 airfoil. The roughness trip heights ranged from 17 to 214% of the undisturbed boundary-layer thickness at the trip location. The chord-based Reynolds number varied from to . This paper shows that trip heights within 50 to 105% of have little influence on the near field and the trailing-edge noise. However, trips with increase the spectra of velocity and pressure fluctuations in the low-frequency range up to 2.5 and 3 dB, respectively, resulting in an increase of the trailing-edge radiated noise up to 4 dB in a similar frequency range. In summary, this paper shows the importance of carefully defining the roughness trip used to trigger a laminar–turbulent transition in aeroacoustic and aerodynamic measurements.

Unsteady interaction of crossflow instability with a forward-facing step

Experiments have been conducted on a swept wing model in a low-turbulence wind tunnel at chord Reynolds number of $2.17 \times 10^{6}$ to investigate the unsteady interaction of a forward-facing step (FFS) with incoming stationary crossflow (CF) vortices. The impact of varying the FFS height on the development and growth of primary and secondary CF disturbances and the ensuing laminar–turbulent transition is quantified through detailed hot-wire anemometry and infrared thermography measurements. The presence of the FFS results in either a critical (i.e. moderate transition advancement) or a supercritical behaviour (i.e. transition advancing abruptly to the FFS location). The arrival of the forced stationary CF vortices at the step is accompanied by their amplification. Unsteady analysis for the critical cases indicates temporal velocity fluctuations following closely the development of the baseline configuration (i.e. agreeing with the development of secondary instabilities). Consequently, laminar breakdown originates from the outer side of the upwelling region of the CF vortices. In contrast, for the supercritical FFS, the laminar breakdown unexpectedly originates from the inner side of the upwelling region. Evidence points to an unsteady mechanism possibly supported by locally enhanced spanwise-modulated shears and the recirculation region downstream of the FFS edge. This mechanism appears to govern the abrupt tripping of the flow in supercritical step cases. The findings in this work provide insight into the unsteady FFS–CF vortex interaction, which is pivotal to understanding the influence of an FFS on the laminar–turbulent boundary-layer transition in swept aerodynamic surfaces.

Quantification technique of transition to turbulence in boundary layers using infrared thermography

The problem of accurate experimental detection of laminar-turbulent transition in three-dimensional boundary layers is discussed. A high-resolving panoramic experimental technique to estimate the onset and length of the transition for aerodynamic applications is proposed and described in detail. The technique includes multistage processing of recorded infrared image sequences to provide a robust detection of the transition in a broad range of base flow parameters. The data processing is fast and virtually hands-free that allows extensive parametric investigation of the transition in experiments. Moreover, an application of statistics procedures to obtain various spanwise-averaged values quantifying different aspects of the transition important for calibrating and validating engineering transition prediction methods is possible.

Modeling of the impact of laminar-turbulent transition on cavitation inception

The main incentive to study cavitation inception has been the accompanying jump of flow-induced noise. The water tunnel tests of cavitation inception have been carried out at Reynolds numbers that are substantially smaller than those which occurs to most of the marine applications. As a result, there are significant scale effects. Employment of trips to enhance laminar-turbulent transition in boundary layers of tested models may seem to be useful to reduce these effects, but the experiment of Arakeri and Acosta showed that the trip use can lead to the trend opposite to the trends observed in flows with the natural transition. This paper gives the theoretical explanation of the difference of trends. The explanation is obtained via computations based on a viscous-inviscid interaction procedure with employing a multi-zone flow model. The viscous part of the flow includes attached boundary layers and viscous separation zones. Possibilities of laminar-turbulent transitions and relaminarization are taken into account, as well as capillarity effect on the surfaces of cavities.

Theoretical, numerical and experimental study of hypersonic boundary layer transition: Blunt circular cone

• Hypersonic boundary layer transition is studied using various approaches. • Leading edge receptivity is crucial for understanding “transition reversal”. • Centreline and crossflow vortices significantly promote transition on a yawed cone. • The newly proposed C - γ - Re θ transition model works well at various flow conditions. • Valuable transition data is acquired in real flight test for a cone model. This article reviews the hypersonic laminar-turbulent boundary layer transition over a blunt circular cone under support of the National Key Research and Development Program of China (contract No. 2016YFA0401200). Various approaches, including theoretical, numerical, experimental and flight test, are combined to obtain a complete perspective of the transition process as well as some fundamental issues for this kind of flow. The interaction between freestream disturbances and the front shock wave is analysed theoretically, and multiple receptivity routes are unveiled using a sub-zone analysis technique. The entire transition process is captured using high-resolution nano-tracer-based planar laser scattering technique in a quiet windtunnel as well as large-scale high-fidelity direct numerical simulations. Using dynamic mode decomposition, the relationship between the second Mach mode and the low-frequency mode in the late stage of the transition is revealed. The “transition reversal” phenomenon is recovered in two of our conventional windtunnels. The numerical results show that leading-edge receptivity varies significantly for different cone bluntness, which might be responsible for the “transition reversal” phenomenon. For the cone at angles of attack, new instabilities including the leeward centerline vortex instability and the crossflow instability, occur and can significantly promote the transition process. A newly developed C - γ - Re θ transition model can give a good transition front prediction compared to the experimental data. The flight test for a cone model conducted within this project is reviewed, and the correlation between the ground experimental data and the flight data is established.

Unsteady Aerodynamic Characteristics of Pitching Flat Plates at Low Reynolds Numbers

A computational study is conducted on thin flat plates to simulate flows of Reynolds numbers at to provide understanding and guidance for micro air vehicles and other low-Reynolds-number airfoil designs. A synergistic effort between experiments and validated and computational fluid dynamics (CFD) tools were used as part of this study. The CFD tool used in this study is an established Reynolds-averaged Navier–Stokes (RANS) solver with a Spalart–Allmaras turbulence model and a correlation-based laminar–turbulent boundary-layer transition model. The computational method was validated against experimental data for flat plates under steady and unsteady kinematic conditions. The objective of the study was to understand unsteady characteristics of a thin flat plate undergoing harmonic pitching (no plunging) around the quarter chord under incompressible flow conditions. Pitching amplitudes were limited to 10 deg to ensure there was no effect of dynamic stall. The focus of this study is to characterize unsteady aerodynamics based on reduced frequency, which was varied between 0.005, 0.05, and 0.5. The unsteady condition of 0.05 was compared against experiments, 2-D RANS, and 3-D hybrid RANS/large-eddy simulation formulations. It was observed that the lift characteristics were reasonably well predicted by the CFD tools when compared to experimental observations. At a reduced frequency of , the pitching motion causes an apparent stabilization of vortices resulting in higher oscillatory lift amplitude than the static value. The 3-D RANS better predicted the pitching-moment characteristics compared to the 2-D RANS, attributed to the breakdown of the leading-edge vortex. It was observed that, although thinner flat-plate airfoils have a higher maximum lift coefficient compared to the thicker NACA 0012, they also produce higher instantaneous pitching moment.

Экспериментальное исследование влияния единичного числа Рейнольдса на положение ламинарно-турбулентного перехода на крыле с дозвуковой передней кромкой при числе Маха 2

Experimental studies of the influence of unit Reynolds number on the laminar-turbulent transition in a supersonic boundary layer of a swept wing with a subsonic leading edge at Mach number 2 are performed. The experiments were performed on a model of a swept wing with a swept angle of the leading edge of 72 degrees and with a 3% profile with a variable chord length in span. The hot-wire measurements showed that a laminar-turbulent transition in a supersonic boundary layer of a swept wing with a subsonic leading edge occurs earlier (~25-30%) than on a model with a supersonic leading edge with the same oncoming flow parameters. It is shown that a change unit Reynolds number insignificant influence the laminar-turbulent transition in the boundary layer of a swept wing with a subsonic leading edge.

Thermal imaging of laminar-turbulent transition on experimental model of swept wing. Creating database of experimental data

The paper is devoted to the problem of creation of experimental database obtained due to systematic extended series of experiments aimed at registering the laminar-turbulent transition in boundary layers on a swept wing model. The experiments were carried out in a low-turbulence wind tunnel in boundary layers developed on ‘suction’ and ‘pressure’ sides of experimental model of an airfoil in a broad range of problem parameters (angle of attack, free-stream velocity, free-stream turbulence etc.). The main feature of performed experiments is using of a new panoramic experimental technique based on application of sensitive infrared camera. Digital postprocessing of obtained thermograms that gave a possibility to detect the position of transition occurred due to amplification of cross-flow instability modes, TS-waves or both of them. The article includes description of the approaches used and the first preliminary results.

Influence of the Entropy Layer on Turbulence Transition of the Supersonic Boundary Layer in the Wake with Isolated Roughnesses

The influence of the entropy layer generated by a small bluntness of the flat plate leading edge on laminar-turbulent supersonic boundary layer transition in the wake behind isolated roughness elements located on the plate surface is studied when the bluntness radius is varied. The experiments are carried out at the Mach number M = 6 over the Reynolds number range (0.4–2.3) × 106. The Reynolds number is calculated on the basis of the free-stream parameters at the point of roughness. The well-known correlation dependences are verified and the transition criterion is estimated.

Statistical Description of Intermittency in the Transition Region at the Low Degree of Flow Turbulence

The statistical method for describing the intermittency coefficient in the region of laminar–turbulent boundary layer transition at the low level of external perturbations is proposed. The method is based on the assumption that turbulence develops when the velocity deviates locally in space and time from the mean by a threshold value. The analytical formulas are obtained for the frequency of occurrence of the coherent structures (turbulent spots and wedges) and for the intermittency coefficient or the mean fraction of the surface area covered by them in the case of two basic scenarios of transition, namely, the enhancement of steady crossflow instability vortices on the swept wing and the Tollmien–Schlichting waves.

On the thresholds of vortex visualisation methods

Since the discovery of vortex structures in turbulence, research in the area of vortex visualisation methods has been of interest. The prevalent Galilean invariant visualisation methods, like Q, Lambda2, and Omega methods, are based on the local pointwise analysis of invariants of the velocity gradient tensor. When plotting the vortex iso-surfaces, different methods will introduce different thresholds for visualisation purpose. Varying the thresholds can greatly change the appearance of the results and the interpretation can be significantly different for the same flow phenomenon. Using our direct numerical simulation data on late stage laminar turbulent boundary layer transition, we will investigate the similarities and differences among several prevailing vortex visualisation methods caused by the selection of thresholds. The sensitivity of thresholds are further explored analytically in a few flow examples. In addition, we propose an appropriate usage of the threshold which will result in remarkably similar looking vortex structures.