Shock Wave/Transitional Boundary-Layer Interactions in Hypersonic Flow

Benay, R.; Chanetz, Bruno; Mangin, Bruno; Perraud, Jean

doi:10.2514/1.10512

Cited by 92 publications

(83 citation statements)

References 10 publications

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“…We shall return to this result after we have reviewed results from DNS. From a comparison of figure 6(a,b) it is possible to identify four different zones of the interaction (as may also be inferred from previous studies such as Benay et al (2006)): (i) the undisturbed flow upstream of the SWBLI where the Stanton number distribution is unaffected by the shock impingement; (ii) a zone with significantly reduced heat transfer, dipping below the laminar boundary-layer solution; (iii) a rapid rise in heat transfer up to a maximum downstream of the shock impingement point; and (iv) a slow recovery of heat-transfer rate towards that of an equilibrium turbulent boundary layer.…”

Section: Flow Visualisation and Mean Stanton Number Distributionsmentioning

confidence: 70%

“…The environment is often of high background noise level, mainly due to sound radiation from turbulent boundary layers on nozzle walls, although some facilities have been designed to be quiet by the use of boundary-layer bleeds (see Schneider 2008 for a review). A study of transition with SWBLI was presented by Benay et al (2006) for a hollow cylinder flare model in the ONERA R2Ch wind tunnel at Mach 5. The stagnation pressure was varied to provide cases that were fully laminar, transitional and fully turbulent.…”

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confidence: 99%

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

et al. 2014

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Strong interactions of shock waves with boundary layers lead to flow separations and enhanced heat transfer rates. When the approaching boundary layer is hypersonic and transitional the problem is particularly challenging and more reliable data is required in order to assess changes in the flow and the surface heat transfer, and to develop simplified models. The present contribution compares results for transitional interactions on a flat plate at Mach 6 from three different experimental facilities using the same instrumented plate insert. The facilities consist of a Ludwieg tube (RWG), an open-jet wind tunnel (H2K) and a high-enthalpy free-piston-driven reflected shock tunnel (HEG). The experimental measurements include shadowgraph and infrared thermography as well as heat transfer and pressure sensors. Direct numerical simulations (DNS) are carried out to compare with selected experimental flow conditions. The combined approach allows an assessment of the effects of unit Reynolds number, disturbance amplitude, shock impingement location and wall cooling. Measures of intermittency are proposed based on wall heat flux, allowing the peak Stanton number in the reattachment regime to be mapped over a range of intermittency states of the approaching boundary layer, with higher overshoots found for transitional interactions compared with fully turbulent interactions. The transition process is found to develop from second (Mack) mode instabilities superimposed on streamwise streaks.

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Section: Flow Visualisation and Mean Stanton Number Distributionsmentioning

confidence: 70%

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confidence: 99%

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

et al. 2014

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“…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).…”

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confidence: 98%

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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“…This result is in agreement with the well-known tendencies relative to the transitional thermal loads, which are larger than the laminar ones and even larger than the turbulent ones. These have been proven by experiments performed at Onera [4] for an identical Reynolds number in natural transition or with triggered transition.…”

Section: Resultsmentioning

confidence: 81%

“…But precisely such interactions give rise to larger heat- §uxes at reattachment, even higher than those obtained in fully turbulent regime. Therefore, this transitional well-documented test case [4] achieved at Onera ¦lls a void in this domain. The third test case completes the range in cold hypersonic since it is relative to a natural fully turbulent shock wave / boundary layer interaction with transition occurring upstream of the separation.…”

Section: Discussionmentioning

confidence: 91%

Low and high enthalpy shock wave / boundary layer interactions around cylinder-flare models

Chanetz

2012

Progress in Flight Physics

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The paper describes four test-cases around the same con¦guration type: a hollow cylinder §are model. The ¦rst three test-cases are relative to cold hypersonic §ow and the fourth one to hot hypersonic §ow: (i) a purely laminar low enthalpy interaction performed at Mach 9.92; (ii) a transitional low enthalpy interaction performed at Mach 5; (iii) a fully low enthalpy turbulent interaction performed at Mach 5; and (iν) a high enthalpy interaction performed at Mach 9.4.

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Shock Wave/Transitional Boundary-Layer Interactions in Hypersonic Flow

Cited by 92 publications

References 10 publications

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

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

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

Low and high enthalpy shock wave / boundary layer interactions around cylinder-flare models

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