Transitional shock-wave/boundary-layer interactions in hypersonic flow

Sandham, Neil D.; Schülein, Erich; Wagner, Alexander; Willems, Sebastian; Steelant, Johan

doi:10.1017/jfm.2014.333

Cited by 117 publications

(78 citation statements)

References 42 publications

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“…This shows 20 that the streak width in the expansion case is larger than on the flat plate. For other wall-normal distances, the increase in the streak width compared to the flat plate case can all be identified.…”

Section: Comparison Of Near Wall Streaks and Two-point Correlationsmentioning

confidence: 76%

“…The code has been previously used for studies of instability, transition and turbulence in high-speed flows [20] . Here we provide a brief summary of the main features of the code and explain the numerical setup for the present study.…”

Section: Numerical Simulation and Inflow Conditionmentioning

confidence: 99%

“…Many DNS databases of wall-bounded supersonic flows are attainable, including fundamental supersonic boundary layer [15][16][17][18] , supersonic boundary layer interaction with shock waves [19,20] , and supersonic boundary layer over a 4 compression ramp [21] etc. Recently Fang et al [22] calculated a supersonic flow at Ma=2.9 over an expansion-compression corner and found that the re-laminarized expansion flow is far from recovered before the separation region in the compression corner.…”

Section: Introductionmentioning

confidence: 99%

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Recovery of a supersonic turbulent boundary layer after an expansion corner

Sun

Sandham

2017

Physics of Fluids

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Supersonic turbulent flows at Mach 2.7 over expansion corners with deflection angles of 0° (flat plate), 2° and 4° have been studied using direct numerical simulation. Distributions of skin friction, pressure, velocity and the boundary layer growth show that the turbulent boundary layer experiences a recovery from a non-equilibrium to an equilibrium state downstream of the expansion corner. Analysis of velocity profiles indicates that the streamwise velocity undergoes a reduction in the near-wall region even though the velocity in the core part of the boundary layer is accelerated after the expansion corner. Growth of the boundary layer was evaluated and a higher shape factor was found in the expansion cases. Turbulence was found to be mostly suppressed downstream of the corner, and throughout the recovery region, even though turbulence is regenerated in the nearwall region. The expansion ramp increases the near wall streak spacing compared to a flat plate and turbulent kinetic energy profiles and budgets exhibit a characteristic two-layer structure. Nearwall turbulence recovers to a balance between the local production and dissipation equilibrium more quickly in the inner layer than in the outer layer. The two-layer structure is due to a history effect of turbulence decay in the outer part of the boundary layer downstream of the expansion corner, with limited momentum and energy exchange between the inner layer and the main stream.

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Section: Comparison Of Near Wall Streaks and Two-point Correlationsmentioning

confidence: 76%

Section: Numerical Simulation and Inflow Conditionmentioning

confidence: 99%

Section: Introductionmentioning

confidence: 99%

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Recovery of a supersonic turbulent boundary layer after an expansion corner

Sun

Sandham

2017

Physics of Fluids

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“…Three different measuring inserts were used for the experiments in H2K (see also Sandham et al 2014). One insert made of PEEK was used for the infrared thermography.…”

Section: Instrumentationmentioning

confidence: 99%

“…Most of the existing publications in this field deal with interactions of shock waves with turbulent boundary layers (Dupont et al 2006; Schülein 2006;Humble et al 2009; Helmer 2011;Grilli et al 2012), some handle the case of interactions of shock waves with laminar boundary layers (Boin et al 2006;Lüdeke and Sandham 2009;Brown and Boyce 2009) but investigations of the interaction between shock waves and transitional boundary layers are rare (Dolling 2001;Arnal and Delery 2004;Benay et al 2006;Vanstone et al 2013). In order to study these phenomena in detail within the framework of the ESA-TRP "laminar to turbulent transition in hypersonic flows," experiments in three different facilities using several measuring techniques have been carried out (Sandham et al 2014). This paper describes the results of the experiments performed in the hypersonic wind tunnel H2K at DLR in Cologne.…”

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confidence: 99%

Experiments on the effect of laminar–turbulent transition on the SWBLI in H2K at Mach 6

2015

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than two and an increase in the heat flux by a factor of more than four. Therefore, big areas of laminar flow are desirable to increase the performance of a propelled vehicle and to reduce the weight of the necessary thermal protection systems of a propelled or re-entry vehicle. On the other hand, turbulent boundary layers have a lower tendency to separate and are preferred in regions of large pressure gradients. Hence, the correct prediction of the transition is essential for the design of future hypersonic vehicles and their thermal protection systems.Shock wave-boundary layer interactions (SWBLI) are inevitable for most of the practical applications. They can cause locally increased pressure and thermal loads, cause separations and have a great influence onto the transition process. Most of the existing publications in this field deal with interactions of shock waves with turbulent boundary layers (Dupont et al. 2006; Schülein 2006;Humble et al. 2009; Helmer 2011;Grilli et al. 2012), some handle the case of interactions of shock waves with laminar boundary layers (Boin et al. 2006;Lüdeke and Sandham 2009;Brown and Boyce 2009) but investigations of the interaction between shock waves and transitional boundary layers are rare (Dolling 2001;Arnal and Delery 2004;Benay et al. 2006;Vanstone et al. 2013). In order to study these phenomena in detail within the framework of the ESA-TRP "laminar to turbulent transition in hypersonic flows," experiments in three different facilities using several measuring techniques have been carried out (Sandham et al. 2014). This paper describes the results of the experiments performed in the hypersonic wind tunnel H2K at DLR in Cologne.The transition process of undisturbed hypersonic boundary layers is most likely driven by instabilities called second (Mack) modes (Mack 1975). These are acoustic waves trapped between the wall and the sonic line. Therefore, Abstract This paper presents the results of the experiments performed in the hypersonic wind tunnel H2K in the framework of the ESA technology research project "laminar to turbulent transition in hypersonic flows". The investigations include the free boundary-layer transition on a flat plate as well as the influence of a shock wave-boundary layer interaction on the transition. The shock is created by a wedge with a small angle of attack resulting in a moderate shock intensity. The experiments were performed at Mach 6.0, at three different unit Reynolds numbers and with a translational displacement of the shock generator. Besides the optical methods-Schlieren photography and infrared thermography-several intrusive sensors were used. High-speed measurements were carried out using PCB and atomic layer thermo pile sensors. Kulite sensors were used for low-and mid-speed pressure measurements. The data analysis includes the comparison of the absolute values, the frequency spectra and wavelets and their distributions in time and space.

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