The Effect of Aspect Ratio on a Mach 2.75 Shock Boundary Layer Interaction Confguration

Galbraith, Daniel; Turner, Mark G.; Orkwis, Paul D.; Weils, Samuel

doi:10.2514/6.2013-861

Cited by 8 publications

(7 citation statements)

References 8 publications

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“…Furthermore, as mentioned in § 1.3, it is widely known that the inclusion of sidewalls can have large three-dimensional effects on the resulting flow (Bruce et al 2011;Burton & Babinsky 2012;Babinsky et al 2013;Benek et al 2013;Galbraith et al 2013;Wang et al 2015). We content that so-called three-dimensional effects, such as pressure disturbances generated by viscous flow phenomena in the Babinsky et al (2013), may also contribute to the 'non-canonical' development of the separation bubble (and hence separation shock) at the tunnel centreline.…”

Section: Free Interaction Theorymentioning

confidence: 84%

“…Matheis & Hickel (2015) investigated transition in SWBLI numerically, but to the best of the authors' knowledge there exists no such experimental study on the topic. Furthermore, Matheis & Hickel (2015) chose not to include the effect of sidewalls; however, it is widely known that the inclusion of sidewalls can have a large effect on the resulting flow (Bruce et al 2011;Burton & Babinsky 2012;Babinsky et al 2013;Benek, Suchyta III & Babinsky 2013;Galbraith et al 2013;Wang et al 2015). An experimental study would naturally include sidewall effects.…”

Section: Motivationmentioning

confidence: 99%

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Confinement effects on regular–irregular transition in shock-wave–boundary-layer interactions

Grossman

Bruce

2018

J. Fluid Mech.

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An oblique shock wave is generated in a Mach 2 flow at a flow deflection angle of 12• . The resulting shock-wave-boundary-layer interaction (SWBLI) at the tunnel wall is observed. A novel traversable shock generator allows the position of the SWBLI to be varied relative to a downstream expansion fan. The relationship between the SWBLI, the expansion fan and the wind tunnel arrangement is studied. Schlieren photography, surface oil flow visualisation, particle image velocimetry and high-spatial-resolution wall pressure measurements are used to investigate the flow. It is observed that stream-normal movement of the shock generator downwards (towards the floor and hence the point of shock reflection) is accompanied by (1) growth in the streamwise extent of the shock-induced boundary layer separation, (2) upstream movement of the shock-induced separation point while the reattachment point remains nearly fixed, (3) an increase in separation shock strength and (4) transition between regular and irregular (Mach) reflection without an increase in incident shock strength. The role of free interaction theory in defining the separation shock angle is considered and shown to be consistent with the present measurements over a short streamwise extent. An SWBLI representation is proposed and reasoned which explains the apparent increase in separation shock strength that occurs without an increase in incident shock strength.

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Section: Free Interaction Theorymentioning

confidence: 84%

Section: Motivationmentioning

confidence: 99%

Confinement effects on regular–irregular transition in shock-wave–boundary-layer interactions

Grossman

Bruce

2018

J. Fluid Mech.

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“…Second, the nozzle may be reconfigured to reach several different desired Mach numbers. The test section in question measures 2.25 wide by 2.75 inches tall [13] and 36 inches long [14]. The converging portion of the nozzle is 12.088 inches long.…”

Section: Physical Wind Tunnelmentioning

confidence: 99%

A Numerical Comparison of Symmetric and Asymmetric Supersonic Wind Tunnels

Clark

2016

46th AIAA Fluid Dynamics Conference

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Supersonic wind tunnels are a vital aspect to the aerospace industry. Both the design and testing processes of different aerospace components often include and depend upon utilization of supersonic test facilities. Engine inlets, wing shapes, and body aerodynamics, to name a few, are aspects of aircraft that are frequently subjected to supersonic conditions in use, and thus often require supersonic wind tunnel testing. There is a need for reliable and repeatable supersonic test

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“…Furthermore, Touber and Sandham (2009) reported that the size of the separation bubble varied with changes in the spanwise extent of the computational domain even when spanwise periodic boundary conditions were employed. Galbraith et al (2013) performed compressible RANS computations of a Mach 2.75 incident shock boundary layer interaction in geometries of aspect ratio one and two. They found significant differences in the amount of blockage and boundary layer separation between the two cases.…”

Section: For Reviews)mentioning

confidence: 99%

Shock boundary layer interactions in a low aspect ratio duct

Campo

Eaton

2015

International Journal of Heat and Fluid Flow

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Experimental data acquired using high resolution two-component particle image velocimetry (PIV) are presented for shock boundary layer interactions (SBLIs) generated by a compression-expansion ramp geometry. The incident oblique shock wave is generated by a sub-boundary layer height ramp inclined 20 • to the M ∞ = 2.05 inflow. Results are presented for two different ramp sizes (h ramp /δ 0 = 0.56 and h ramp /δ 0 = 0.93), and compared to a previously documented h ramp /δ 0 = 0.20 case. For each case, mean velocitiy and turbulent statistics for both the SBLI at the foot of the compression ramp and the incident/reflected SBLI on the opposite wall are analyzed. Data are acquired in several streamwise-vertical planes across the span of the low aspect ratio duct in order to document spanwise non-uniformities and confinement effects, with the specific goal of producing a high quality experimental database for CFD validation. The angles of the incident and reflected shock waves become steeper as the side wall is approached, due to the lower velocities within the side wall boundary layer. Mean flow reversal is observed near the spanwise centerline of the duct, but only instantaneous flow reversal is seen closer to the side walls. These spanwise non-uniformities are more prominent for the stronger interactions caused by the larger ramp geometry. For this case there is no nominally two-dimensional region of the flow, and a Mach stem occurs in the core of the flow, causing a significant subsonic wake. The shock excursion length scale relative to the incoming boundary layer thickness, L ex /δ 0 , is measured for all of the shock features and found to be significantly lower than values reported in the literature for similar flows. Furthermore, L ex /δ 0 for the reflected shock does not depend on the strength of the incident shock or the size of the separated zone.

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The Effect of Aspect Ratio on a Mach 2.75 Shock Boundary Layer Interaction Confguration

Cited by 8 publications

References 8 publications

Confinement effects on regular–irregular transition in shock-wave–boundary-layer interactions

Confinement effects on regular–irregular transition in shock-wave–boundary-layer interactions

A Numerical Comparison of Symmetric and Asymmetric Supersonic Wind Tunnels

Shock boundary layer interactions in a low aspect ratio duct

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