Stability of preseparation boundary layer on the leading edge of a thin airfoil

1982 ◽  
Vol 17 (6) ◽  
pp. 860-867 ◽  
Author(s):  
A. I. Ruban
Keyword(s):  
Boundary Layer ◽  
Leading Edge ◽  
Thin Airfoil

INFLUENCE OF THE PLATE LEADING-EDGE SHAPE AND THICKNESS ON THE BOUNDARY LAYER LAMINAR-TURBULENT TRANSITION IN HYPERSONIC FLOW

2019 ◽  
Vol 50 (5) ◽  
pp. 461-481
Author(s):  
Sergei Vasilyevich Aleksandrov ◽  
Evgeniya Andreevna Aleksandrova ◽  
Volf Ya. Borovoy ◽  
Andrey Vyacheslavovich Gubernatenko ◽  
Vladimir Evguenyevich Mosharov ◽  
...  
Keyword(s):  
Boundary Layer ◽  
Hypersonic Flow ◽  
Leading Edge ◽  
Turbulent Transition ◽  
Laminar Turbulent Transition

Characteristics of a separated flow past a semicircular leading-edge airfoil model under different imposed pressure gradient

2021 ◽  
pp. 095441002110212
Author(s):  
K Anand ◽  
KT Ganesh
Keyword(s):  
Boundary Layer ◽  
Particle Image Velocimetry ◽  
Pressure Gradient ◽  
Particle Image ◽  
Separation Bubble ◽  
Separated Flow ◽  
Leading Edge ◽  
Field Measurements ◽  
Bench Mark ◽  
Image Velocimetry

The effect of pressure gradient on a separated boundary layer past the leading edge of an airfoil model is studied experimentally using electronically scanned pressure (ESP) and particle image velocimetry (PIV) for a Reynolds number ( Re) of 25,000, based on leading-edge diameter ( D). The features of the boundary layer in the region of separation and its development past the reattachment location are examined for three cases of β (−30°, 0°, and +30°). The bubble parameters such as the onset of separation and transition and the reattachment location are identified from the averaged data obtained from pressure and velocity measurements. Surface pressure measurements obtained from ESP show a surge in wall static pressure for β = −30° (flap deflected up), while it goes down for β = +30° (flap deflected down) compared to the fundamental case, β = 0°. Particle image velocimetry results show that the roll up of the shear layer past the onset of separation is early for β = +30°, owing to higher amplification of background disturbances compared to β = 0° and −30°. Downstream to transition location, the instantaneous field measurements reveal a stretched, disoriented, and at instances bigger vortices for β = +30°, whereas a regular, periodically shed vortices, keeping their identity past the reattachment location, is observed for β = 0° and −30°. Above all, this study presents a new insight on the features of a separation bubble receiving a disturbance from the downstream end of the model, and these results may serve as a bench mark for future studies over an airfoil under similar environment.

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 Effects of Tip Leakage Flow on the Aerodynamic Performance and Stability of an Axial-Flow Transonic Compressor Stage

Energies ◽  
2021 ◽  
Vol 14 (14) ◽  
pp. 4168
Author(s):  
Botao Zhang ◽  
Xiaochen Mao ◽  
Xiaoxiong Wu ◽  
Bo Liu
Keyword(s):  
Boundary Layer ◽  
Shock Wave ◽  
Axial Flow ◽  
Leading Edge ◽  
Rotating Stall ◽  
Tip Clearance ◽  
Leakage Flow ◽  
Tip Leakage Flow ◽  
Tip Leakage ◽  
Transonic Compressor

To explain the effect of tip leakage flow on the performance of an axial-flow transonic compressor, the compressors with different rotor tip clearances were studied numerically. The results show that as the rotor tip clearance increases, the leakage flow intensity is increased, the shock wave position is moved backward, and the interaction between the tip leakage vortex and shock wave is intensified, while that between the boundary layer and shock wave is weakened. Most of all, the stall mechanisms of the compressors with varying rotor tip clearances are different. The clearance leakage flow is the main cause of the rotating stall under large rotor tip clearance. However, the stall form for the compressor with half of the designed tip clearance is caused by the joint action of the rotor tip stall caused by the leakage flow spillage at the blade leading edge and the whole blade span stall caused by the separation of the boundary layer of the rotor and the stator passage. Within the investigated varied range, when the rotor tip clearance size is half of the design, the compressor performance is improved best, and the peak efficiency and stall margin are increased by 0.2% and 3.5%, respectively.

scholarly journals Leading Edge Blowing to Mimic and Enhance the Serration Effects for Aerofoil

2021 ◽  
Vol 11 (6) ◽  
pp. 2593
Author(s):  
Yasir Al-Okbi ◽  
Tze Pei Chong ◽  
Oksana Stalnov
Keyword(s):  
Boundary Layer ◽  
Angle Of Attack ◽  
Leading Edge ◽  
Boundary Layer Separation ◽  
Shear Instability ◽  
Vortical Flow ◽  
High Angle ◽  
High Angle Of Attack ◽  
Layer Separation ◽  
Aerodynamic Performances

Leading edge serration is now a well-established and effective passive control device for the reduction of turbulence–leading edge interaction noise, and for the suppression of boundary layer separation at high angle of attack. It is envisaged that leading edge blowing could produce the same mechanisms as those produced by a serrated leading edge to enhance the aeroacoustics and aerodynamic performances of aerofoil. Aeroacoustically, injection of mass airflow from the leading edge (against the incoming turbulent flow) can be an effective mechanism to decrease the turbulence intensity, and/or alter the stagnation point. According to classical theory on the aerofoil leading edge noise, there is a potential for the leading edge blowing to reduce the level of turbulence–leading edge interaction noise radiation. Aerodynamically, after the mixing between the injected air and the incoming flow, a shear instability is likely to be triggered owing to the different flow directions. The resulting vortical flow will then propagate along the main flow direction across the aerofoil surface. These vortical flows generated indirectly owing to the leading edge blowing could also be effective to mitigate boundary layer separation at high angle of attack. The objectives of this paper are to validate these hypotheses, and combine the serration and blowing together on the leading edge to harvest further improvement on the aeroacoustics and aerodynamic performances. Results presented in this paper strongly indicate that leading edge blowing, which is an active flow control method, can indeed mimic and even enhance the bio-inspired leading edge serration effectively.

Swept wing boundary-layer receptivity to localized surface roughness

2012 ◽  
Vol 711 ◽  
pp. 516-544 ◽  
Author(s):  
David Tempelmann ◽  
Lars-Uve Schrader ◽  
Ardeshir Hanifi ◽  
Luca Brandt ◽  
Dan S. Henningson
Keyword(s):  
Boundary Layer ◽  
Surface Roughness ◽  
Practical Interest ◽  
Leading Edge ◽  
Constant Factor ◽  
Computationally Efficient ◽  
Swept Wing ◽  
Predictive Tool ◽  
Free Stream Turbulence ◽  
Streamwise Location

AbstractThe receptivity to localized surface roughness of a swept-wing boundary layer is studied by direct numerical simulation (DNS) and computations using the parabolized stability equations (PSEs). The DNS is laid out to reproduce wind tunnel experiments performed by Saric and coworkers, where micron-sized cylinders were used to trigger steady crossflow modes. The amplitudes of the roughness-induced fundamental crossflow wave and its superharmonics obtained from nonlinear PSE solutions agree excellently with the DNS results. A receptivity model using the direct and adjoint PSEs is shown to provide reliable predictions of the receptivity to roughness cylinders of different heights and chordwise locations. Being robust and computationally efficient, the model is well suited as a predictive tool of receptivity in flows of practical interest. The crossflow mode amplitudes obtained based on both DNS and PSE methods are 40 % of those measured in the experiments. Additional comparisons between experimental and PSE data for various disturbance wavelengths reveal that the measured disturbance amplitudes are consistently larger than those predicted by the PSE-based receptivity model by a nearly constant factor. Supplementary DNS and PSE computations suggest that possible natural leading-edge roughness and free-stream turbulence in the experiments are unlikely to account for this discrepancy. It is more likely that experimental uncertainties in the streamwise location of the roughness array and cylinder height are responsible for the additional receptivity observed in the experiments.

Characterization of Shock Wave–Endwall Boundary Layer Interactions in a Transonic Compressor Rotor

1988 ◽  
Vol 110 (3) ◽  
pp. 386-392 ◽  
Author(s):  
D. C. Rabe ◽  
A. J. Wennerstrom ◽  
W. F. O’Brien
Keyword(s):  
Boundary Layer ◽  
Shock Wave ◽  
Air Force ◽  
Leading Edge ◽  
Transonic Compressor ◽  
Laser Measurements ◽  
Compressor Rotor ◽  
Boundary Layer Interaction ◽  
Air Force Base

The passage shock wave–endwall boundary layer interaction in a transonic compressor was investigated with a laser transit anemometer. The transonic compressor used in this investigation was developed by the General Electric Company under contract to the Air Force. The compressor testing was conducted in the Compressor Research Facility at Wright-Patterson Air Force Base, OH. Laser measurements were made in two blade passages at seven axial locations from 10 percent of the axial blade chord in front of the leading edge to 30 percent of the axial blade chord into the blade passage. At three of these axial locations, laser traverses were taken at different radial immersions. A total of 27 different locations were traversed circumferentially. The measurements reveal that the endwall boundary layer in this region is separated from the core flow by what appears to be a shear layer where the passage shock wave and all ordered flow seem to end abruptly.

Keel Design for Low Viscous Drag

1987 ◽  
Author(s):  
Clifford J. Obara ◽  
C. P. van Dam
Keyword(s):  
Boundary Layer ◽  
Laminar Flow ◽  
Laminar Boundary Layer ◽  
Leading Edge ◽  
Boundary Layer Transition ◽  
Viscous Drag ◽  
Hydrodynamic Characteristics ◽  
Sweep Angle ◽  
Layer Transition ◽  
Layer Flow

In this paper, foil and planform parameters which govern the level of viscous drag produced by the keel of a sailing yacht are discussed. It is shown that the application of laminar boundary-Layer flow offers great potential for increased boat speed resulting from the reduction in viscous drag. Three foil shapes have been designed and it is shown that their hydro­dynamic characteristics are very much dependent on location and mode of boundary-Layer transition. The planform parameter which strongly affects the capabilities of the keel to achieve laminar flow is lea ding-edge sweep angle. The two significant phenomena related to keel sweep angle which can cause premature transition of the laminar boundary layer are crossflow instability and turbulent contamination of the leading-edge attachment line. These flow phenomena and methods to control them are discussed in detail. The remaining factors that affect the maintainability of laminar flow include surface roughness, surface waviness, and freestream turbulence. Recommended limits for these factors are given to insure achievability of laminar flow on the keel. In addition, the application of a simple trailing-edge flap to improve the hydrodynamic characteristics of a foil at moderate-to-high leeway angles is studied.

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