Visualization of High-Speed Boundary-Layer Transition with FPA Infrared Technique

Muller, Lilac; Henckels, A.

doi:10.1007/978-3-322-86573-1_31

Cited by 3 publications

(7 citation statements)

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“…It is based on a thermal energy balance over finite volume elements and assumes a semi-infinite wall thickness (Müller and Henckels 1997;Henckels and Gruhn 2004). The recovery factor for the post-processing of the infrared images is always set to the value of a laminar boundary layer r = √ Pr.…”

Section: Stanton Number From Infrared Picturesmentioning

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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Section: Stanton Number From Infrared Picturesmentioning

confidence: 99%

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

2015

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“…The preliminary computations have been performed on a model very similar to the one of interest (the geometry is slightly different from the CIRA model, being the cone angle equal to 2.9 degrees instead of 3) and supposed to have a smooth surface. The results were compared with the available literature data [15]. In [15] it is shown that, on this kind of geometry, transition generally needs a finite length to fully develop.…”

Section: A Cfd Analysis Descriptionmentioning

confidence: 99%

“…The results were compared with the available literature data [15]. In [15] it is shown that, on this kind of geometry, transition generally needs a finite length to fully develop. Fig.…”

Section: A Cfd Analysis Descriptionmentioning

confidence: 99%

“…3 reports the experimental transition points at M=6 on a cone geometry for several Reynolds numbers and shows that the transition length change as the nose radius and Reynolds number is changing. However, the bluntness paradox is clearly visible on the transition onset( [15]).…”

Section: A Cfd Analysis Descriptionmentioning

confidence: 99%

“…The total number of points necessary to construct a d th -order model in k independent variables that predicts the response within a specified tolerance is given as ( [30] and [32]): (15) Here, σ is the standard deviation in the measurement, λ is the precision requirement for the response surface, and t α and t β are statistics related to acceptable inference error probabilities for Type I and Type II errors, respectively.…”

Section: Angle Of Attackmentioning

confidence: 99%

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Hypersonic Laminar-Turbulent Transition Experiment Design: From Wind Tunnel Model Definition to MDOE Approach

Fauci

Donelli

Schettino

2010

48th AIAA Aerospace Sciences Meeting Including the New Horizons Forum and Aerospace Exposition

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This paper describes the Design of Experiment to study the laminar-turbulence transition phenomenon in Hypersonic regime on a 3-deg half-angle sphere-cone model. The huge number of factors (as bluntness, distributed roughness, Mach and Reynolds numbers, etc.), that affect the laminar-turbulent transition, makes this phenomenon very complex and expensive to study with a typical experimental approach. As consequence, a Modern Design Of Experiment approach was adopted to develop highly efficient experiment designs yielding results to within a specified accuracy and, at the same time, saving resources in terms of costs and time consuming. Preliminary numerical analyses have been performed on several sphere-cone geometries for several values of the Mach and Reynolds numbers in order to define the basic requirements for the experimental test campaign. Moreover, an innovative and simple rough nose realization methodology definition was an essential step of the experiment design phase, as it conferred to MDOE technique, the mandatory freedom in terms of number of noses and values of roughness selection. Nomenclature CIRA = Italian Aerospace Research Centre TPS = thermal protection system MDOE = Modern Design of Experiment OFAT = One Factor At Time k = Roughness height inducing transition δ = Boundary layer thickness 1 Researcher/Test Engineer, Transonic Wind Tunnel.Laboratory, a.marino@cira.it

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Instability and Transition Measurements in the Mach-6 Quiet Tunnel

Wheaton

Juliano

Berridge

et al. 2009

39th AIAA Fluid Dynamics Conference

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The Boeing/AFOSR Mach-6 Quiet Tunnel achieved quiet flow to a stagnation pressure of 163 psia in Dec. 2008, the highest value observed so far. It remains quiet at pressures above 160 psia. Under noisy conditions, nozzle-wall boundary-layer separation and the associated tunnel shutdown appear to propagate slowly upstream, whereas under quiet conditions, the propagation is very rapid. A new diffuser insert has been designed, fabricated, and installed in the tunnel in order to start larger blunt models and increase run time. A flared cone with a circular-arc geometry was designed to generate large second-mode N factors under quiet flow conditions. When the computed N factor was 13, large instability waves were measured under quiet flow conditions using fast pressure sensors, but the flow remained laminar. Transition was observed only under noisy conditions. A laminar instability was detected in the wake of an isolated roughness element in the boundary layer on the nozzle wall; this appears to be the first such measurement at hypersonic speeds.

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Visualization of High-Speed Boundary-Layer Transition with FPA Infrared Technique

Cited by 3 publications

References 4 publications

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

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

Hypersonic Laminar-Turbulent Transition Experiment Design: From Wind Tunnel Model Definition to MDOE Approach

Instability and Transition Measurements in the Mach-6 Quiet Tunnel

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