1205 Generation Mechanism of Noise Induced by Flow Separation around the Leading Edge

Abstract The present study focuses on the acoustics of a turbocharger centrifugal compressor from a spark-ignition internal combustion engine. Whoosh noise is typically the primary concern for this type of compressor, which is loosely characterized by broadband sound elevation in the 4 to 13 kHz range. To identify the generation mechanism of broadband whoosh noise, the present study combines three approaches: three-dimensional (3D) computational fluid dynamics (CFD) predictions, experiments, and modal decomposition of 3D CFD results. After establishing the accuracy of predictions, flow structures and time-resolved pressures are closely examined in the vicinity of the main blade leading edge. This reveals the presence of rotating instabilities that may interact with the rotor blades to generate noise. An azimuthal modal decomposition is performed on the predicted pressure field to determine the number of cells and the frequency content of these rotating instabilities. The strength of the rotating instabilities and the frequency range in which noise is generated as a consequence of the rotor-rotating instability interaction, is found to correspond well with the qualitative trend of the whoosh noise that is measured several duct diameters upstream of the rotor blades. The variation of whoosh frequency range between low and high rotational speeds is interpreted through this analysis. It is also found that the whoosh noise primarily propagates along the duct as acoustic azimuthal modes. Hence, the inlet duct diameter, which governs the cut-off frequency for multi-dimensional acoustic modes, determines the lower frequency bound of the broadband noise.

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Flow Separation Control of Backward-Facing Step Airfoil NACA0015 by Blowing Technique

Diyala Journal of Engineering Sciences ◽

10.24237/djes.2019.12111 ◽

2019 ◽

Vol 12 (1) ◽

pp. 99-119

Author(s):

Khuder N. Abed

Keyword(s):

Flow Separation ◽

Numerical Study ◽

Leading Edge ◽

Distribution Coefficients ◽

Open Circuit ◽

Experimental Investigations ◽

Backward Facing Step ◽

Ansys Fluent ◽

Flow Separation Control ◽

Naca 0015

The aim of this paper is to control the flow separation above backward-facing step (BFS) airfoil type NACA 0015 by blowing method. The flow field over airfoil has been studied both experimentally and computationally. The study was divided into two parts: a practical study through which NACA 0015 type with a backward -facing step (located at 44.4% c from leading edge) on the upper surface containing blowing holes parallel to the airfoil chord was used. The tests were done over two-dimensional airfoil in an open circuit suction subsonic wind tunnel with flow velocity 25m/s to obtain the pressure distribution coefficients. A numerical study was done by using ANSYS Fluent software version 16.0 on three models of NACA 0015, the first one has backward-facing step without blowing, the second with single blowing holes and the third have multi blowing holes technique. Both studies (experimental and numerical) were done at low Reynolds number (Re=4.4x105) and all models have chord length 0.27m.The experimental investigations and CFD simulations have been performed on the same geometry dimensions, it has been observed that the flow separation on the airfoil can be delayed by using velocity blowing (30m/s) on the upper surface. The multi blowing holes with velocity improved the aerodynamics properties.The multi blowing holes and single blowing hole thesame effect onpressure distribution coefficients

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Leading-edge Flow Separation

Separation of Flow ◽

10.1016/b978-0-08-013441-3.50013-7 ◽

1970 ◽

pp. 452-530 ◽

Cited By ~ 1

Author(s):

PAUL K. CHANG

Keyword(s):

Flow Separation ◽

Leading Edge ◽

Edge Flow

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Study of three types of wave leading edge on the performance of industrial turbine blade cascade

Aircraft Engineering and Aerospace Technology ◽

10.1108/aeat-06-2020-0115 ◽

2020 ◽

Vol ahead-of-print (ahead-of-print) ◽

Author(s):

Hamed Ghandi ◽

Reza Aghaei Togh ◽

Abolghasem Mesgarpoor Tousi

Keyword(s):

Turbine Blade ◽

Flow Separation ◽

Numerical Solutions ◽

Incidence Angle ◽

Leading Edge ◽

Content Type ◽

Passive Flow Control ◽

Passive Flow ◽

Geometrical Features ◽

Control Methodology

Purpose The blade profile and its geometrical features play an important role in the separation of the boundary layer on the blade. Modifying the blade geometry, which might lead to the delay or elimination of the flow separation, can be considered as a passive flow control methodology. This study aims to find a novel and inexpensive way to reduce loss with appropriate modifications on the leading edge of the turbine blade. Design/methodology/approach Three types of wave leading edges were designed with different wavelengths and amplitudes. The selected numbers for the wave characteristics were based on the best results of previous studies. Models with appropriate and independent meshing have been simulated and studied by a commercial software. The distribution of the loss at different planes and mid-plane velocity vectors were shown. The mass flow average of loss at different incidence angles was calculated for the reference blade and modified ones for the sake of comparison. Findings The results show that in all three types of modified blades compared to the reference blade, the elimination of flow separation is observed and therefore the reduction of loss at the critical incidence angle of I = –15°. As the amplitude of the wave increased, the amount of loss growing up, while the increase in wavelength caused the loss to decrease. Originality/value The results of the present numerical analysis were validated by the laboratory results of the reference blade. The experimental study of modified blades can be used to quantify numerical solutions.

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Leading-Edge Surface-Manipulated Flow Separation from an Airfoil

Journal of Aircraft ◽

10.2514/1.36998 ◽

2008 ◽

Vol 45 (6) ◽

pp. 2171-2173 ◽

Cited By ~ 1

Author(s):

Hong J. Zhang ◽

Yu Zhou

Keyword(s):

Flow Separation ◽

Leading Edge ◽

Edge Surface

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High Mach Number Leading-Edge Flow Separation Control Using AC DBD Plasma Actuators

50th AIAA Aerospace Sciences Meeting including the New Horizons Forum and Aerospace Exposition ◽

10.2514/6.2012-906 ◽

2012 ◽

Cited By ~ 23

Author(s):

Christopher Kelley ◽

Patrick Bowles ◽

John Cooney ◽

Chuan He ◽

Thomas Corke ◽

...

Keyword(s):

Mach Number ◽

Flow Separation ◽

Leading Edge ◽

Plasma Actuators ◽

Separation Control ◽

Dbd Plasma ◽

Flow Separation Control ◽

High Mach Number ◽

High Mach ◽

Edge Flow

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Investigations of the Flow Through a High Pressure Ratio Centrifugal Impeller

Volume 5: Turbo Expo 2002, Parts A and B ◽

10.1115/gt2002-30394 ◽

2002 ◽

Cited By ~ 17

Author(s):

Gernot Eisenlohr ◽

Hartmut Krain ◽

Franz-Arno Richter ◽

Valentin Tiede

Keyword(s):

High Pressure ◽

Flow Separation ◽

Pressure Ratio ◽

Leading Edge ◽

Centrifugal Impeller ◽

Angle Distribution ◽

Blade Angle ◽

Minor Variation ◽

Design Point ◽

High Pressure Ratio

In an industrial research project of German and Swiss Turbo Compressor manufacturers a high pressure ratio centrifugal impeller was designed and investigated. Performance measurements and extensive laser measurements (L2F) of the flow field upstream, along the blade passage and downstream of the impeller have been carried out. In addition to that, 3D calculations have been performed, mainly for the design point. Results have been presented by Krain et al., 1995 and 1998, Eisenlohr et al., 1998 and Hah et al.,1999. During the design period of this impeller a radial blade at the inlet region was mandatory to avoid a rub at the shroud due to stress reasons. The measurements and the 3D calculations performed later, however, showed a flow separation at the hub near the leading edge due to too high incidence. Additionally a rather large exit width and a high shroud curvature near the exit caused a flow separation near the exit, which is enlarged by the radially transported wake of the already addressed hub separation. Changes to the hub blade angle distribution to reduce the hub incidence and an adaptation of the shroud blade angle distribution for the same impeller mass-flow at the design point were investigated by means of 3D calculations first with the same contours at hub and shroud; this was followed by calculations with a major change of the shroud contour including an exit width change with a minor variation of the hub contour. These calculations showed encouraging results; some of them will be presented in conjunction with the geometry data of the original impeller design.

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Film Cooling on Highly Loaded Blades With Main Flow Separation—Part I: Heat Transfer

Journal of Turbomachinery ◽

10.1115/1.4006568 ◽

2012 ◽

Vol 135 (1) ◽

Cited By ~ 1

Author(s):

Reinaldo A. Gomes ◽

Reinhard Niehuis

Keyword(s):

Heat Transfer ◽

Flow Separation ◽

Film Cooling ◽

Separation Bubble ◽

Leading Edge ◽

Local Heat Transfer ◽

Local Heat ◽

Main Flow ◽

Cooling Effectiveness ◽

Film Cooling Effectiveness

Film cooling experiments were run at the high speed cascade wind tunnel of the University of the Federal Armed Forces Munich. The investigations were carried out with a linear cascade of highly loaded turbine blades. The main objectives of the tests were to assess the film cooling effectiveness and the heat transfer in zones with main flow separation. Therefore, the blades were designed to force the flow to detach on the pressure side shortly downstream of the leading edge and reattach at about half of the axial chord. In this zone, film cooling rows are placed among others for a reduction of the size of the separation bubble. The analyzed region on the blade is critical due to the high heat transfer present at the leading edge and at the reattachment line after the main flow separation. Film cooling can contribute to a reduction of the size of the separation bubble reducing aerodynamic losses, however, in general, it increases heat transfer due to turbulent mixing. The reduction of the size of the separation bubble might also be twofold, since it acts like a thermal insulator on the blade and reducing the size of the bubble might lead to a stronger heating of the blade. Film cooling should, therefore, take both into account: first, a proper protection of the surface and second, reducing aerodynamic losses, diminishing the extension of the main flow separation. While experimental results of the adiabatic film cooling effectiveness were shown in previous publications, the local heat transfer is analyzed in this paper. Emphasis is also placed upon analyzing, in detail, the flow separation process. Furthermore, the tests comprise the analysis of the effect of different outlet Mach and Reynolds numbers and film cooling. In part two of this paper, the overall film cooling effectiveness is addressed. Local heat transfer is still difficult to predict with modern numerical tools and this is especially true for complex flows with flow separation. Some numerical results with the Reynolds averaged Navier-Stokes (RANS) and large eddy simulation (LES) show the capability of a commercial solver in predicting the heat transfer.

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