Transition effect on aerodynamic performance of compressor profile

2020 ◽  
Vol 92 (4) ◽  
pp. 611-620
Author(s):  
Ryszard Szwaba ◽  
Piotr Kaczyński ◽  
Piotr Doerffer
Keyword(s):  
Boundary Layer ◽  
Shock Wave ◽  
Aerodynamic Performance ◽  
Negative Influence ◽  
Good Choice ◽  
Boundary Layer Transition ◽  
Roughness Height ◽  
Compressor Cascade ◽  
Layer Transition ◽  
Content Type

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.

Numerical study of boundary layer transition using intermittency model

2019 ◽  
Vol 91 (8) ◽  
pp. 1156-1168 ◽  
Author(s):  
Massoud Tatar ◽  
Mojtaba Tahani ◽  
Mehran Masdari
Keyword(s):  
Boundary Layer ◽  
Shear Stress ◽  
Transport Equation ◽  
Transition Point ◽  
Static Pressure ◽  
Boundary Layer Transition ◽  
Layer Transition ◽  
Content Type ◽  
Reduced Frequency ◽  
Pressure Suction

Purpose In this paper, the applicability of shear stress transport k-ω model along with the intermittency concept has been investigated over pitching airfoils to capture the laminar separation bubble (LSB) position and the boundary layer transition movement. The effect of reduced frequency of oscillations on boundary layer response is also examined. Design/methodology/approach A two-dimensional computational fluid dynamic code was developed to compute the effects of unsteadiness on LSB formation, transition point movement, pressure distribution and lift force over an oscillating airfoil using transport equation of intermittency accompanied by the k-ω model. Findings The results indicate that increasing the angle of attack over the stationary airfoil causes the LSB size to shorten, leading to a rise in wall shear stress and pressure suction peak. In unsteady cases, both three- and four-equation models are capable of capturing the experimentally measured transition point well. The transition is delayed for an unsteady boundary layer in comparison with that for a static airfoil at the same angle of attack. Increasing the unsteadiness of flow, i.e. reduced frequency, moves the transition point toward the trailing edge of the airfoil. This increment also results in lower static pressure suction peak and hence lower lift produced by the airfoil. It was also found that the fully turbulent k-ω shear–stress transport (SST) model cannot capture the so-called figure-of-eight region in lift coefficient and the employment of intermittency transport equation is essential. Practical implications Boundary layer transition and unsteady flow characteristics owing to airfoil motion are both important for many engineering applications including micro air vehicles as well as helicopter blade, wind turbine and aircraft maneuvers. In this paper, the accuracy of transition modeling based on intermittency transport concept and the response of boundary layer to unsteadiness are investigated. Originality/value As a conclusion, the contribution of this paper is to assess the ability of intermittency transport models to predict LSB and transition point movements, static pressure distribution and aerodynamic lift variations and boundary layer flow pattern over dynamic pitching airfoils with regard to oscillation frequency effects for engineering problems.

Numerical Simulation of the Boundary Layer Transition in Turbomachinery Flows

2001 ◽  
Author(s):  
Hans Thermann ◽  
Michael Müller ◽  
Reinhard Niehuis
Keyword(s):  
Boundary Layer ◽  
Separated Flow ◽  
Suction Side ◽  
Boundary Layer Transition ◽  
Test Cases ◽  
Compressor Cascade ◽  
Layer Transition ◽  
Transonic Compressor ◽  
Transition Models ◽  
Transition Modeling

The objective of the presented work is to investigate models which simulate boundary layer transition in turbomachinery flows. This study focuses on separated-flow transition. Computations with different algebraic transition models are performed three-dimensionally using an implicit Navier-Stokes flow solver. Two different test cases have been chosen for this investigation: First, a linear transonic compressor cascade, and second an annular subsonic compressor cascade. Both test cases show three-dimensional flow structures with large separations at the side-walls. Additionally, laminar separation bubbles can be observed on the suction and pressure side of the blades of the annular subsonic cascade whereas a shock-induced separation can be found on the suction side of the blades of the linear transonic cascade. Computational results are compared with experiments and the effect of transition modeling is analyzed. It is shown that the prediction of the boundary layer development can be substantially improved compared to fully turbulent computations when algebraic transition models are applied.

Scramjet Flow and Intake SBLI: Technical Challenges and Case Study

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.

Effect of Wake Negative Jet Flow on the Boundary Layer Transition of Compressor Cascade

2018 ◽  
Author(s):  
Yangang Wang ◽  
Qijie Shao ◽  
Wenbing Hu
Keyword(s):  
Boundary Layer ◽  
Reynolds Number ◽  
Numerical Study ◽  
Jet Flow ◽  
Boundary Layer Transition ◽  
Turbulent Intensity ◽  
Compressor Cascade ◽  
Suction Surface ◽  
Layer Transition ◽  
Transition Behavior

The present paper performed a numerical study on a high-loaded and high turning compressor cascade, where the unsteady boundary layer transition behavior on the cascade blade undergoing negative jet flow is revealed. The two-equation SST turbulence model coupled with Langtry-Menter transition model is verified and applied on all the computations in present study. Reynolds number and turbulent intensity are selected as two dominate candidates which can significantly influence the transition behavior and their effect were examined. Results show that under all the tested case (i.e., varying Reynolds number and turbulence intensity), the flow structures on the suction surface of the blade are rolled up when the unsteady negative jet flow directly impacting on the blade. However, the unsteady wake from upstream has not influenced the boundary layer. For high Reynolds number (i.e., Re = 400,000) the rolling up and shed of the boundary layer only occurs at blade trailing edge. The wake is evidenced be able to bring more energy into the boundary layer and thus separation and loss can be significantly decayed and reduced. Moreover, decreasing the turbulent intensity would in practical decay the transition in the boundary layer and therefore make the boundary layer easy to separate.

Experimental Investigation of the Three-Dimensional Flow in an Annular Compressor Cascade

1988 ◽  
Vol 110 (4) ◽  
pp. 467-478 ◽  
Author(s):  
H. D. Schulz ◽  
H. D. Gallus
Keyword(s):  
Boundary Layer ◽  
Experimental Investigation ◽  
Three Dimensional ◽  
Suction Side ◽  
Boundary Layer Transition ◽  
Compressor Cascade ◽  
Three Dimensional Flow ◽  
Layer Transition ◽  
Integral Boundary ◽  
Dimensional Flow

A detailed experimental investigation was carried out to examine the influence of blade loading on the three-dimensional flow in an annular compressor cascade. Data were acquired over a range of incidence angles. Included are airfoil and endwall flow visualization, measurement of the static pressure distribution on the flow passage surfaces, and radial-circumferential traverse measurements. The data indicate the formation of a strong vortex near the rear of the blade passage. This vortex transports low-momentum fluid close to the hub toward the blade suction side and seems to be partly responsible for the occurrence of a hub corner stall. The effect of increased loading on the growth of the hub corner stall and its impact on the passage blockage are discussed. Detailed mapping of the blade boundary layer was done to determine the loci of boundary layer transition and flow separation. The data have been compared with results from an integral boundary layer method.

scholarly journals Experimental Investigation of Boundary Layer Transition and Turbulence Structures on a Highly Loaded Compressor Cascade

1995 ◽  
Author(s):  
Dirk Wunderwald ◽  
Leonhard Fottner
Keyword(s):  
Boundary Layer ◽  
Suction Side ◽  
Boundary Layer Transition ◽  
Reynolds Stresses ◽  
Flow Conditions ◽  
Compressor Cascade ◽  
Layer Transition ◽  
Engine Operation ◽  
Inlet Flow ◽  
Free Stream Turbulence

Detailed measurements have been performed on a compressor cascade in order to obtain information about the overall performance, the state of the boundary layer, and the topology of turbulent boundary layers. The analysis of profile pressure distributions and wake traverse measurements across the midspan section of the cascade blade provide information on the loss behaviour. Using surface-mounted hot-film gauges on the suction side of the measuring blade different transition phenomena have been investigated under the influence of various inlet flow conditions representative of engine operation. Extensive measurements with 3D-hot-sensor anemometry have been evaluated to show essential features of the turbulent boundary layer. The results point out the dependence of turbulence characteristics, e.g. turbulent kinetic energy distribution and Reynolds stresses, on the inlet flow conditions and the upstream boundary layer development. The influence of free-stream turbulence intensity is discussed and the non-isotropy of the Reynolds normal stresses is presented.

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