CFD Study of Effects of Boundary Layer Suction on Transonic SC(2)-0714 Airfoil Performance

2017 ◽  
Author(s):  
Ruslan Khamedov ◽  
Ruslan Baitlessov ◽  
Luis Rojas-Solórzano
Keyword(s):  
Boundary Layer ◽  
Shock Waves ◽  
Isotropic Turbulence ◽  
Free Stream ◽  
Computational Study ◽  
Leading Edge ◽  
Design Stage ◽  
Stream Velocity ◽  
Local Boundary ◽  
Boundary Layer Suction

The complete understanding of the aerodynamics of wings and blades under transonic conditions represents a substantial challenge in the design of modern airplanes and turbomachinery. Transonic flow over airfoils may result in appearance of shock waves, which lead to increase in drag if not properly considered during the design stage. Therefore, it is a major challenge to design transonic airfoils such that potential appearance of shock waves is foreseen and negative drag effects are minimized. This paper presents the computational study of the SC(2)-0714 airfoil, focusing on its aerodynamics characteristics at Reynolds number of 35 × 106 and angle of attack of 2 and 10 degrees which are the most common operational conditions of transonic wings using this airfoil. The study was undertaken at free-stream Mach 0.72. The numerical simulation was conducted using the finite volume method on platform ANSYS CFX™ and solving the Reynolds-Averaged Navier-Stokes, mass conservation and energy equations. Mesh verification and model validation are presented. The latter is developed by using two different isotropic turbulence models: k-ω and Shear Stress Transport (SST) and the comparison of results with NASA experimental data to determine the best among the treated models. Thereafter, effects of local boundary-layer suction on shock wave strength and characteristics during transonic speed are analyzed for the two aforementioned angles of attack. Two suction slots were placed along the airfoil contour to determine their control effectiveness when compared to standard closed-contour airfoil. Suction slots were placed at the leading edge and in the middle of the upper camber of the airfoil with inflow in the normal direction to the surface. The slot length was 2.5 % of the chord with inflow velocity of 30%, 40% and 50% of free-stream velocity. Effects of suction slots were assessed on the wake region and by computing the resulting lift-to-drag ratio. Concluding remarks on the turbulence model and global aerodynamics performance of the airfoil are presented.

Numerical simulation of a spatially developing accelerating boundary layer over roughness

2015 ◽  
Vol 780 ◽  
pp. 192-214 ◽  
Author(s):  
J. Yuan ◽  
U. Piomelli
Keyword(s):  
Numerical Simulation ◽  
Boundary Layer ◽  
Friction Coefficient ◽  
Isotropic Turbulence ◽  
Free Stream ◽  
Rough Wall ◽  
Stream Velocity ◽  
Mean Velocity ◽  
Wake Field ◽  
Near Wall

The direct numerical simulation of an accelerating boundary layer over a rough wall has been carried out to investigate the coupling between the effects of roughness and strong free-stream acceleration. While the favourable pressure gradient is sufficient to achieve quasi-laminarization on a smooth wall, the flow reversion is prevented on a rough wall, and a higher friction coefficient, a faster increase of turbulence intensity compared to the free-stream velocity and more isotropic turbulence near the wall are observed. The logarithmic region of the mean-velocity profile presents an initial decrease in slope as in the smooth case, but soon recovers, as the fully rough regime is reached and a new overlap region is established. A strong coupling between the roughness and acceleration effects develops as roughness leads to more responsive turbulence and prevents the strong acceleration from stabilizing the turbulence, and the acceleration intensifies the velocity scale of the wake field (i.e. the near-wall spatial heterogeneity of the time-averaged velocity distribution). The combined effect is a ‘rougher’ surface as the flow accelerates. In addition, the link between the local values of the free stream and the near-wall velocity depends on the flow history; this explains the different flow responses observed in previous studies, in terms of friction coefficient, turbulent kinetic energy and Reynolds-stress anisotropy. This study elucidates the near-wall flow dynamics, which may be used to explain other non-canonical flows over rough walls.

Linear stability of a gas boundary layer flowing past a thin liquid film over a flat plate

2001 ◽  
Vol 436 ◽  
pp. 321-352 ◽  
Author(s):  
NIKOLAOS A. PELEKASIS ◽  
JOHN A. TSAMOPOULOS
Keyword(s):  
Boundary Layer ◽  
Liquid Film ◽  
Flat Plate ◽  
Free Stream ◽  
Leading Edge ◽  
Critical Value ◽  
Stream Velocity ◽  
Absolute Instability ◽  
Interfacial Waves ◽  
The Family

The flow of a gas stream past a flat plate under the influence of rainfall is investigated. As raindrops sediment on the flat plate, they coalesce to form a water film that flows under the action of shear from the surrounding gas stream. In the limit of (a) large Reynolds number, Re, in the gas phase, (b) small rainfall rate, r˙, compared to the free-stream velocity, U∞, and (c) small film thickness compared to the thickness of the boundary layer that surrounds it, a similarity solution is obtained that predicts growth of the liquid film like x3/4; x denotes dimensionless distance from the leading edge. The flow in the gas stream closely resembles the Blasius solution, whereas viscous dissipation dominates inside the film. Local linear stability analysis is performed, assuming nearly parallel base flow in the two streams, and operating in the triple-deck regime. Two distinct families of eigenvalues are identified, one corresponding to the well-known Tollmien–Schlichting (TS) waves that originate in the gas stream, and the other corresponding to an interfacial instability. It is shown that, for the air–water system, the TS waves are convectively unstable whereas the interfacial waves exhibit a pocket of absolute instability, at the streamwise location of the applied disturbance. Moreover, it is found that as the inverse Weber number (We−1) increases, indicating the increasing effect of surface tension compared to inertia, the pocket of absolute instability is translated towards larger distances from the leading edge and the growth rate of unstable waves decreases, until a critical value is reached, We−1 ≈ We−1c, beyond which the family of interfacial waves becomes convectively unstable. Increasing the inverse Froude number (Fr−1), indicating the increasing effect of gravity compared to inertia, results in the pocket of absolute instability shrinking until a critical value is reached, Fr−1 ≈ Fr−1c, beyond which the family of interfacial waves becomes convectively unstable. As We−1 and Fr−1 are further increased, interfacial waves are eventually stabilized, as expected. In this context, increasing the rainfall rate or the free-stream velocity results in extending the region of absolute instability over most of the airfoil surface. Owing to this behaviour it is conjectured that a global mode that interacts with the boundary layer may arise at the interface and, eventually, lead to three-dimensional waves (rivulets), or, under extreme conditions, even premature separation.

Unsteady boundary-layer transition in low-pressure turbines

2011 ◽  
Vol 681 ◽  
pp. 370-410 ◽  
Author(s):  
JOHN D. COULL ◽  
HOWARD P. HODSON
Keyword(s):  
Boundary Layer ◽  
Shear Layer ◽  
Flat Plate ◽  
Free Stream ◽  
Separation Bubble ◽  
Leading Edge ◽  
Transition Process ◽  
Free Stream Velocity ◽  
Stream Velocity ◽  
Free Stream Turbulence

This paper examines the transition process in a boundary layer similar to that present over the suction surfaces of aero-engine low-pressure (LP) turbine blades. This transition process is of significant practical interest since the behaviour of this boundary layer largely determines the overall efficiency of the LP turbine. Modern ‘high-lift’ blade designs typically feature a closed laminar separation bubble on the aft portion of the suction surface. The size of this bubble and hence the inefficiency it generates is controlled by the transition between laminar and turbulent flow in the boundary layer and separated shear layer. The transition process is complicated by the inherent unsteadiness of the multi-stage machine: the wakes shed by one blade row convect through the downstream blade passages, periodically disturbing the boundary layers. As a consequence, the transition to turbulence is multi-modal by nature, being promoted by periodic and turbulent fluctuations in the free stream and the inherent instabilities of the boundary layer. Despite many studies examining the flow behaviour, the detailed physics of the unsteady transition phenomena are not yet fully understood. The boundary-layer transition process has been studied experimentally on a flat plate. The opposing test-section wall was curved to impose a streamwise pressure distribution typical of modern high-lift LP turbines over the flat plate. The presence of an upstream blade row has been simulated by a set of moving bars, which shed wakes across the test section inlet. Further upstream, a grid has been installed to elevate the free-stream turbulence to a level believed to be representative of multi-stage LP turbines. Extensive particle imaging velocimetry (PIV) measurements have been performed on the flat-plate boundary layer to examine the flow behaviour. In the absence of the incoming bar wakes, the grid-generated free-stream turbulence induces relatively weak Klebanoff streaks in the boundary layer which are evident as streamwise streaks of low-velocity fluid. Transition is promoted by the streaks and by the inherent inflectional (Kelvin–Helmholtz (KH)) instability of the separation bubble. In unsteady flow, the incoming bar wakes generate stronger Klebanoff streaks as they pass over the leading edge, which convect downstream at a fraction of the free-stream velocity and spread in the streamwise direction. The region of amplified streaks convects in a similar manner to a classical turbulent spot: the leading and trailing edges travel at around 88% and 50% of the free-stream velocity, respectively. The strongest disturbances travel at around 70% of the free-stream velocity. The wakes induce a second type of disturbance as they pass over the separation bubble, in the form of short-span KH structures. Both the streaks and the KH structures contribute to the early wake-induced transition. The KH structures are similar to those observed in the simulation of separated flow transition with high free-stream turbulence by McAuliffe & Yaras (ASME J. Turbomach., vol. 132, no. 1, 2010, 011004), who observed that these structures originated from localised instabilities of the shear layer induced by Klebanoff streaks. In the current measurements, KH structures are frequently observed directly under the path of the wake. The wake-amplified Klebanoff streaks cannot affect the generation of these structures since they do not arrive at the bubble until later in the wake cycle. Rather, the KH structures arise from an interaction between the flow disturbances in the wake and localised instabilities in the shear layer, which are caused by the weak Klebanoff streaks induced by the grid turbulence. The breakdown of the KH structures to small-scale turbulence occurs a short time after the wake has passed over the bubble, and is largely driven by the arrival of the wake-amplified Klebanoff streaks from the leading edge. During this process, the re-attachment location moves rapidly upstream. The minimum length of the bubble occurs when the strongest wake-amplified Klebanoff streaks arrive from the leading edge; these structures travel at around 70% of the free-stream velocity. The bubble remains shorter than its steady-flow length until the trailing edge of the wake-amplified Klebanoff streaks, travelling at 50% of the free-stream velocity, convect past. After this time, the reattachment location moves aft on the surface as a consequence of a calmed flow region which follows behind the wake-induced turbulence.

Experimental study of heat transfer in the boundary layer of a flat plate with the impact of weak shock waves on the leading edge

2020 ◽  
Author(s):  
V. L. Kocharin ◽  
A. A. Yatskikh ◽  
D. S. Prishchepova ◽  
A. V. Panina ◽  
Yu. G. Yermolaev ◽  
...  
Keyword(s):  
Heat Transfer ◽  
Boundary Layer ◽  
Experimental Study ◽  
Shock Waves ◽  
Flat Plate ◽  
Leading Edge ◽  
Weak Shock ◽  
The Impact

scholarly journals Numerical Analysis of Effect of Leading-Edge Rotating Cylinder on NACA0021 Symmetric Airfoil

2019 ◽  
Vol 4 (7) ◽  
pp. 11-17
Author(s):  
Md. Abdus Salam ◽  
Vikram Deshpande ◽  
Nafiz Ahmed Khan ◽  
M. A. Taher Ali
Keyword(s):  
Free Stream ◽  
Numerical Study ◽  
Tangential Velocity ◽  
Leading Edge ◽  
Rotating Cylinder ◽  
Aerodynamic Performance ◽  
Stream Velocity ◽  
Surface Boundary ◽  
Lower Velocity ◽  
Numerical Studies

The moving surface boundary control (MSBC) has been a Centre stage study for last 2-3 decades. The preliminary aim of the study was to ascertain whether the concept can improve the airfoil characteristics. Number of experimental and numerical studies pointed out that the MSBC can superiorly enhance the airfoil performance albeit for higher velocity ratios (i.e. cylinder tangential velocity to free stream velocity). Although abundant research has been undertaken in this area on different airfoil performances but no attempt was seen to study effect of MSBC on NACA0021 airfoil for and also effects of lower velocity ratios. Thus, present paper focusses on numerical study of modified NACA 0021 airfoil with leading edge rotating cylinder for velocity ratios (i.e.) between 1 to 1.78 at different angles of attack. The numerical study indicates that the modified airfoil possess better aerodynamic performance than the base airfoil even at lower velocity ratios (i.e. for velocity ratios 0.356 and beyond). The study also focusses on reason for improvement in aerodynamic performance by close look at various parameters.

scholarly journals Experimental study of weak shock waves influence on the supersonic boundary layer of the flat plate model

2019 ◽  
Vol 196 ◽  
pp. 00018 ◽  
Author(s):  
Vasiliy Kocharin ◽  
Aleksandr Kosinov ◽  
Yuriy Yermolayev ◽  
Nikolay Semionov
Keyword(s):  
Boundary Layer ◽  
Experimental Study ◽  
Shock Waves ◽  
Flat Plate ◽  
Leading Edge ◽  
Weak Shock ◽  
Supersonic Boundary Layer ◽  
Complex Flow ◽  
Complex Flow Structure ◽  
Constant Temperature Anemometer

The experimental study of the effect of weak shock waves on the supersonic boundary layer of the flat plate with a blunt leading edge (the radius of bluntness was r = 2.5 mm) with Mach number M = 2.5 and zero angle of attack was carried out. The measurements were carried out using the constant temperature anemometer. The paper presents a complex flow structure on the surface of the model. High-intensity peaks were found in the regions of the disturbed flow. Also the spectral analysis of perturbations was performed. It is found that the supersonic boundary layer on a flat plate is very sensitive to the effect of weak shock waves.

Eddy Viscosity Distributions in Compressible Turbulent Boundary Layers with Injection

1971 ◽  
Vol 22 (2) ◽  
pp. 169-182 ◽  
Author(s):  
L. C. Squire
Keyword(s):  
Boundary Layer ◽  
Eddy Viscosity ◽  
Free Stream ◽  
Turbulent Boundary Layers ◽  
Stream Velocity ◽  
Mixing Length ◽  
Carbon Dioxide Injection ◽  
Displacement Thickness ◽  
Porous Surfaces ◽  
Injection Ratio

SummaryShear stress, eddy viscosity and mixing length distributions have been obtained from measured boundary-layer developments over porous surfaces with air and carbon dioxide injection at Mach numbers up to M=3·6. It is found that, if the eddy viscosity is non-dimensionalised by dividing by the product of the free-stream velocity and the kinematic displacement thickness then this non-dimensional ratio is almost independent of injection ratio, but decreases slightly with Mach number.

scholarly journals Prediction of Heat Transfer From Laminar Boundary Layers, With Emphasis on Large Free-Stream Velocity Gradients and Highly Cooled Walls

1966 ◽  
Vol 88 (3) ◽  
pp. 249-256 ◽  
Author(s):  
L. H. Back ◽  
A. B. Witte
Keyword(s):  
Heat Transfer ◽  
Boundary Layer ◽  
Shear Stress ◽  
Velocity Gradient ◽  
Free Stream ◽  
Variable Density ◽  
Similarity Solutions ◽  
Free Stream Velocity ◽  
Stream Velocity ◽  
Gradient Parameter

Laminar boundary-layer heat transfer and shear-stress predictions from existing similarity solutions are extended in an approximate way to perfect gas flows with a large free-stream velocity gradient parameter β and variable density-viscosity product ρμ across the boundary layer resulting from a highly cooled wall. The dimensionless enthalpy gradient at the wall gw′, to which the heat flux is related, is found not to vary appreciably with β. Thus the application of similarity solutions on a local basis to predict heat transfer from accelerated flows to an arbitrary surface may be a reasonable approximation involving a minimum amount of calculation time. Unlike gw′, the dimensionless velocity gradient at the wall fw″, to which the shear stress is related, is strongly dependent on β.

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