“…Figure 1(c) shows a range of SWBLI types occurring within a supersonic compression inlet (e.g. figure 1a,b), which is one of the most commonly studied applications (Longley & Greitzer 1992;Smart 2001;Ogawa & Babinsky 2006;Babinsky & Ogawa 2008;Babinsky, Oorebeek & Cottingham 2013;Loth et al 2013).…”
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.
“…Figure 1(c) shows a range of SWBLI types occurring within a supersonic compression inlet (e.g. figure 1a,b), which is one of the most commonly studied applications (Longley & Greitzer 1992;Smart 2001;Ogawa & Babinsky 2006;Babinsky & Ogawa 2008;Babinsky, Oorebeek & Cottingham 2013;Loth et al 2013).…”
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.
“…The diffuser inlet height was 17.3δ which expanded to 20.7δ at the data extraction plane, denoted as the “Measuring Plane” (MP) and this was located 43.7δ downstream of the inlet ( x =58.7δ). The basis of the diffuser design was performed in a previous study ( 31 ) .Figure 3. Schematic of the ramped vane: (a) key dimensions in terms of δ and (b) height of the three ramped vanes.Figure 4.…”
Section: Computational Methodologymentioning
confidence: 99%
“…The diffuser inlet height was 17.3δ which expanded to 20.7δ at the data extraction plane, denoted as the "Measuring Plane" (MP) and this was located 43.7δ downstream of the inlet (x = 58.7δ). The basis of the diffuser design was performed in a previous study (31) .…”
Section: Computational Domain Boundary Conditions and Grid Resolutionmentioning
confidence: 99%
“…The reduction in the turbulent kinetic energy distribution was discussed, as well. The flowfield employed to investigate such effects is a normal shock boundary layer interaction followed by a subsequent subsonic diffuser, which is a flow that has been shown to have the critical features associated with supersonic inlets ( 31 ) .Figure 1. Schematics of primary flow features: (a) no flow control device and RVs with a height of (b) 0.38δ, (c) 0.57δ and (d) 0.75δ.…”
A novel vortex generator design positioned upstream of a normal shock followed by a subsequent diffuser was investigated using large eddy simulations. In particular, “ramped-vane” flow control devices with three different heights relative to the incoming boundary layer thickness (0.34δ, 0.52δ and 0·75δ) were placed in a supersonic boundary layer with a freestream Mach number of 1.3 and a Reynolds number of 2400 based on the momentum thickness. This is the first numerical study to investigate the size effect of the ramped-vane for flow control device in terms of shape factor, flow separation and flow unsteadiness. The results showed that these devices generated strong streamwise vortices that entrained high-momentum fluid to the near-wall region and increased turbulent mixing. The devices also decreased shock-induced flow separation, which resulted in a higher downstream skin friction in the diffuser. In general, the largest ramped-vane (0.75δ) produced the largest reductions in flow separation, shape factor and overall unsteadiness. These results and a careful review of the literature study also determined the quantitative correlation of optimum VG height with Mach number, wherebyh/δ~1 is often optimum for incompressible flows while higher Mach numbers lead to small optimum heights, tending towards h/δ~0.45 atM=2.5.
“…Other recent studies have been carried out on unsteadiness of SBLI [26,27] and control methods such as bleed, vortex generator, and micro actuators [28][29][30]. Also, some recent experiments and theoretical e orts have been made to investigate the e ect of ow control methods such as bleed on SBLI for supersonic and hypersonic inlets [31][32][33][34][35]. The main idea of these notable papers is to discuss the most general aspects of shock wave boundary-layer interactions relevant to supersonic air inlets.…”
Abstract. Experiments were conducted to study various kinds of Shock wave/BoundaryLayer Interaction (SBLI) in an axisymmetric mixed-compression inlet. Experimental ndings were compared and veri ed by numerical solutions where possible. Di erent types of SBLI relevant to the mixed-compression inlets are classi ed. Interactions of normal shock wave/boundary-layer at subcritical condition and in buzz condition are investigated using Schlieren and shadowgraph ow visualization as well as unsteady pressure recordings. The data is compared with the CFD predictions. Interactions of cowl lip re ected oblique shock and the terminal normal shock with the spike boundary-layer at both critical and supercritical operations, which lead to pseudo-shock phenomena, are also studied. Experimental pressure recordings are used for validation and discussion. For near-critical throttling values, interaction of internal compression oblique shocks with boundary-layer and pseudo-shock phenomenon is dominant. Formation of lambda shock due to interaction of separated boundary-layer with normal shock wave is investigated. Each type of ow interaction phenomena has di erent e ects on the stability and performance of the inlet. Interaction of terminal normal shock with internal duct boundary-layer causes pseudoshock phenomenon that increases ow distortion and reduces total pressure recovery. Interaction of normal shock with cone boundary-layer causes buzz instability and degrades inlet performance.
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