Comparison of techniques for determination of boundary layer transition in shock wave induced flows

Purpose The purpose of this paper is to study experimentally the effect of transition and also the roughness height on the flow structure of the shock wave boundary layer interaction in the blades passage of a compressor cascade. Design/methodology/approach A model of a turbine compressor passage was designed and assembled in a transonic wind tunnel. In the experiment, the distributed roughness with different heights and locations was used to induce transition upstream of the shock wave. Findings Recommendation regarding the roughness parameters for the application depends on what is more important as goal, whether the reduction of losses or unsteadiness. In case if more important are the losses reduction, a good choice for the roughness location seems to be the one close to the shock wave position. Research limitations/implications The knowledge gained by this paper will enable the implementation of an effective laminar flow technology for engines in which the interaction of a laminar boundary layer with a shock wave takes place in the propulsion system and causes severe problems. Originality/value The paper focuses on the influence of the boundary layer transition induced by different roughness values and locations on aerodynamic performance of a compressor cascade. Very valuable results were obtained in the roughness application for the boundary layer transition control, demonstrating a positive effect in changing the nature of the interaction and also some negative influence in case of oversized roughness height, which cannot be found in the existing literature.

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Experimental determination of the three-dimensional vorticity field in the boundary-layer transition process

Journal of Fluid Mechanics ◽

10.1017/s0022112084002603 ◽

1984 ◽

Vol 149 (-1) ◽

pp. 179 ◽

Cited By ~ 36

Author(s):

D. R. Williams ◽

H. Fasel ◽

F. R. Hama

Keyword(s):

Boundary Layer ◽

Experimental Determination ◽

Three Dimensional ◽

Transition Process ◽

Boundary Layer Transition ◽

Vorticity Field ◽

Layer Transition

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Report No. 637, Determination of boundary-layer transition on three symmetrical airfoils in the N. A. C. A. full-scale wind tunnel

Journal of the Franklin Institute ◽

10.1016/s0016-0032(39)90856-2 ◽

1939 ◽

Vol 227 (6) ◽

pp. 866-869

Keyword(s):

Boundary Layer ◽

Wind Tunnel ◽

Full Scale ◽

Boundary Layer Transition ◽

Layer Transition

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Scramjet Flow and Intake SBLI: Technical Challenges and Case Study

Applied Mechanics and Materials ◽

10.4028/www.scientific.net/amm.315.344 ◽

2013 ◽

Vol 315 ◽

pp. 344-348

Author(s):

Yu Feng Yao

Keyword(s):

Boundary Layer ◽

Shock Wave ◽

Model Development ◽

Basic Research ◽

Boundary Layer Transition ◽

Layer Transition ◽

Wave Boundary Layer ◽

Research Perspectives ◽

Wave Boundary

This paper reviews some basic research areas associated with Scramjet-powered hypersonic flying vehicle, particularly the forebody boundary-layer transition and intake shock-wave boundary-layer interactions (SBLI). Some technical and physical challenges in aerodynamics, aero-thermodynamics, aero-design are visited with focuses being placed on hypersonic boundary-layer transition process and its underlying physical mechanics, feasible physics-based engineering transition prediction methods, and physics-based modelling of shock-shock, shock-wave/boundary-layer interactions of Scramjet flows. Experimental, analytical and numerical studies of previously relevant studies have also been summarized with a total of twelve transition/intake configurations that can be used as benchmarks for validating physical model development and numerical simulation tools. A case study of Scramjet intake SBLI has been carried out by using computational fluid dynamics approach to understand shock induced flow separation and its consequent influences on combustion performance, along with research perspectives discussed accordingly.

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Transition Effect on the Flow Dynamics in a Compressor Blade Passage

Shock and Vibration ◽

10.1155/2021/5586177 ◽

2021 ◽

Vol 2021 ◽

pp. 1-9

Author(s):

Ryszard Szwaba ◽

Piotr Kaczynski ◽

Janusz Telega

Keyword(s):

Boundary Layer ◽

Shock Wave ◽

Flow Control ◽

Flow Dynamics ◽

Boundary Layer Transition ◽

The Other ◽

Control Methods ◽

Layer Transition ◽

Blade Passage ◽

Transition Control

An experimental investigation was carried out to study the effect of the boundary layer transition on the flow dynamics in the blade passage of a compressor cascade. A model of a turbine compressor passage was designed and assembled in a transonic wind tunnel for this purpose. Two different flow control methods were used in the experiment to induce the transition upstream of the shock wave, one concerning the microstep application and the other using distributed roughness strips. Two locations of spanwise microsteps for the transition trigger were chosen, one at the leading edge and the other closer to the shock wave position. The distributed roughness in the form of standard sandpaper strips with different heights was applied in three various locations on the blade upstream of the shock. The main objective of investigations is to present the influence of the laminar and turbulent shock wave-boundary layer interaction on the flow dynamics in a compressor fan passage, and the specific objective is to show how the parameters of particular transition control methods affect the flow dynamics in the investigated channel. The major challenge for this research was to minimize the disturbance caused by the microstep or roughness to the laminar boundary layer, while still ensuring a successful transition. Very interesting results were obtained in the flow control application for the boundary layer transition control, demonstrating a positive effect on the flow unsteadiness in changing the nature of the interaction.

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