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H Babinsky - One of the best experts on this subject based on the ideXlab platform.
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effect of reynolds number on a Normal Shock Wave transitional boundary layer interaction over a curved surface
Experiments in Fluids, 2019Co-Authors: A Coschignano, H Babinsky, Nicholas R Atkins, J SernaAbstract:The interaction between a Normal Shock Wave and a boundary layer is investigated over a curved surface for a Reynolds number range, based on boundary-layer growing length x, of $$0.44\times 10^6\le \text {Re}_x\le 1.09\times 10^6$$. The upstream boundary layer develops around the leading edge of the model before encountering a $$M$$ $$\sim $$1.4 Normal Shock. This is followed by adverse pressure gradients. The Shock position and strength are kept constant as $$\text {Re}$$ is progressively varied. Infra-red thermography is used to determine the nature of the upstream boundary layer. Across the $$\text {Re}$$ range, this is observed to vary from fully laminar to fully turbulent across the entire span. Regardless of the boundary-layer state, the interaction remains benign in nature, without large scale Shock-induced separation or unsteadiness. Schlieren images show a pronounced oblique Wave developing upstream of the main Shock for the laminar cases, this is believed to correspond to the separation and subsequent transition of the laminar shear layer. Downstream of the Shock, in the presence of adverse pressure gradients, the boundary-layer growth rate is inversely proportional to $$\text {Re}$$. Nonetheless, across the entire range of inflow conditions the boundary layer recovers quickly to a healthy turbulent boundary layer. This suggests the upstream boundary-layer state, and its transition mechanism, to have little effect on the outcome of its interaction with a Normal Shock Wave.
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vortex generators for corner separation caused by Shock Wave boundary layer interactions
Journal of Aircraft, 2019Co-Authors: Shunsuke Koike, H BabinskyAbstract:Wind tunnel experiments were conducted to investigate the effect of vortex generators on a transonic corner flow separation, resulting from the interaction of a Normal Shock Wave with a turbulent b...
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Normal Shock Wave turbulent boundary layer interactions in transonic intakes at incidence
2018 AIAA Aerospace Sciences Meeting, 2018Co-Authors: A Coschignano, H Babinsky, C Sheaf, E PlattAbstract:© 2018, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. The flow field around a transonic engine inlet lip at high incidence is investigated for a variety of flow conditions around the design point. Generally, the flow on the upper surface of the lip is characterised by a supersonic region, terminated by a near-Normal Shock Wave. At the nominal design point, the Shock is not strong enough to cause significant flow separation, resulting only in marginal losses in pressure recovery. Off-design conditions were explored by altering the angle of attack as well as changing the mass flow rate over the upper lip, intended to mimic the effect of an increase in engine flow. The results suggest that angle of attack has the greatest effect on the flow field. In particular, even a relatively small increase of 2 ◦ can lead to large and highly unsteady flow separation with an associated Shock oscillation. Both qualitative and quantitative measurements suggest a noticeably reduced aerodynamic performance resulting from higher incidence operation. In contrast, an increase of up to 5.2% in mass flow over the upper part of the intake lip did not result in large separated regions or flow-field unsteadiness.
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influence of transition on the flow downstream of Normal Shock Wave boundary layer interactions
54th AIAA Aerospace Sciences Meeting, 2016Co-Authors: Tsc Davidson, H BabinskyAbstract:© 2016, Todd Davidson and Holger Babinsky. This paper reports recent work undertaken to investigate the interaction between Normal Shocks and laminar/transitional boundary layers. Laminar flow is desirable for its low skin friction, but problems are expected with Shock interactions as the boundary layer will be prone to separation. Experiments have been performed on a flat plate to explore the fundamental interaction without complex geometries and flow fields; and on a transonic aerofoil to research the influence of the imposed pressure field. The boundary layer state does not appear to significantly affect the downstream flow development, contrary to expectations, as the laminar separation is so small. On the aerofoil, a similar result is observed, with no change in the buffet onset Mach number and little change in the wake. This implies that Normal Shock Wave–laminar boundary layer interactions are not as detrimental as feared, and are in fact quite benign.
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transition location effects on Normal Shock Wave boundary layer interactions
53rd AIAA Aerospace Sciences Meeting, 2015Co-Authors: Tsc Davidson, H BabinskyAbstract:© 2015 by Todd Davidson and Holger Babinsky. Published by the American Institute of Aeronautics and Astronautics, Inc. This paper reports recent work undertaken to investigate the interaction between Normal Shocks and laminar/transitional boundary layers. Laminar ow is desirable for its low skin friction, but problems are expected with Shock interactions as the boundary layer will be prone to separation. Experiments have been performed on a at plate to explore the fundamental interaction without complex geometries and ow fields, and on a transonic aerofoil to research the in uence of the imposed pressure field. The boundary layer state before the Shock on the at plate does not appear to significantly affect the downstream ow development, contrary to expectations, as the laminar separation is so small. On the aerofoil, a similar result is observed, with no change in the buffet onset Mach number. This implies that Normal Shock Wave-laminar boundary layer interactions are not as detrimental as originally feared, and are in fact quite benign.
Kazuyasu Matsuo - One of the best experts on this subject based on the ideXlab platform.
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three dimensional Normal Shock Wave boundary layer interaction in a rectangular duct
AIAA Journal, 2005Co-Authors: Taro Handa, Mitsuharu Masuda, Kazuyasu MatsuoAbstract:The three-dimensional flow structure induced by Normal Shock-Wave/turbulent boundary-layer interaction in a constant-area rectangular duct is investigated by a laser-induced-fluorescence method. This diagnostic system uses an argon-ion laser as a light source, and the target gas is dry nitrogen with iodine seeded as a fluorescence material. The Mach-number distributions in the duct are obtained from the measured fluorescence intensity, and the threedimensional flow pattern in the expansion region downstream of the initial Shock Wave is clarified. In addition to this, the region having locally higher Mach number near the duct corners is observed immediately behind the Shock Wave, and the three-dimensional shape of the boundary layers is found. These flow characteristics are reproduced by solving the Navier‐Stokes equations numerically. The calculated result reveals that the complicated Shock-Wave configuration is formed at the duct corner because of the interaction of two bifurcated Shock Waves developed on the two perpendicularly adjacent walls. The simple flow model is also constructed by considering this interaction. This model can explain very well the three-dimensional flow characteristics.
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experimental investigation on the three dimensional structure of Normal Shock Wave boundary layer interaction in a constant area rectangular duct
Transactions of the Japan Society of Mechanical Engineers. B, 2004Co-Authors: Taro Handa, Mitsuharu Masuda, Kazuyasu MatsuoAbstract:The three-dimensional flow structure induced by Normal Shock Wave/turbulent boundary-layer interaction in a constant area rectangular duct is investigated by a laser-induced fluorescence method. This diagnostic system uses an argon-ion laser as a light source, and the target gas is dry nitrogen with iodine seeded as a fluorescence material.The Mach number distributions in the duct are obtained from the measured fluorescence intensity, and the three-dimensional flow pattern in the expansion region downstream of the initial Shock Wave is clarified. In addition to this, the region having locally higher Mach number at the duct corners is observed just behind the Shock Wave, and the three-dimensional shape of the boundary layers induced by the Shock Wave is found. These flow characteristics are explained with the simple flow model constructed by considering the interaction between two bifurcated Shock Waves developed on the two perpendicularly adjacent walls.
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three dimensional structure of Normal Shock Wave turbulent boundary layer interaction in a rectangular duct by a laser induced fluorescence method
Engineering sciences reports Kyushu University, 2002Co-Authors: Taro Handa, Mitsuharu Masuda, Kazuyasu MatsuoAbstract:The three-dimensional flow structure induced by Normal Shock Wave/turbulent boundary-layer interaction in a constant area rectangular duct is investigated by a laser-induced fluorescence method. This diagnostic system uses an argon-ion laser with iodine seeded as fluorescence material. The Mach number distributions in the duct are obtained, and the structure of the flow field is clarified including the three-dimensional pattern of the boundary-layer separation induced by a Shock Wave.
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spanwise pressure measurements of weak Normal Shock Wave turbulent boundary layer interactions in a supersonic nozzle
Engineering sciences reports Kyushu University, 1995Co-Authors: Jong Woo Hong, Yoshiaki Miyazato, Kazuyasu MatsuoAbstract:The characteristics of flow field along spanwise direction induced by the interaction of weak Normal Shock Wave with turbulent boundary layer in a supersonic nozzle were investigated experimentally. In the range of just upstream Shock Mach number about 1.10 to 1.80, detailed time mean wall static pressure measurements were carried out along spanwise direction using multiple pressure transducers. Also, the relations between variations of upstream boundary layer and behavior of interaction flow fields were investigated. As a result, it is revealed that upstream boundary layer thickness influences the behavior of interaction somewhat for case of well-separated flow. And, another one is revealed that there is some spanwise pressure difference in unseparated flow, however, no so much difference is found in well-separated flow case.
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Normal Shock Wave oscillations in supersonic diffusers
Shock Waves, 1993Co-Authors: Kazuyasu Matsuo, Heuy Dong KimAbstract:The present paper describes experimental investigations for Shock oscillations caused by Normal Shock Wave/turbulent boundary layer interaction in a supersonic diffuser. An array of wall-mounted transducers and especially a line image sensor for the nonintrusive detection of Shock displacements were employed to investigate the interactions at low supersonic speeds. The line image sensor was collimated with a conventional schlieren optical system and was a good indicative of capturing the Shock oscillating motions in the present configuration. This study shows that the amplitude of the Shock motions increases with approaching flow Mach number, and the cause of oscillation of the Shock Wave can, however, be independent of the Mach number. In addition, the present system employed to determine the Shock Wave positions and displacements can be effectively applied to a variety of practical problems.
Mortaza Mani - One of the best experts on this subject based on the ideXlab platform.
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numerical investigation of a Normal Shock Wave boundary layer interaction in a 4 3 aspect ratio test section
54th AIAA Aerospace Sciences Meeting, 2016Co-Authors: Miranda P Pizzella, Sally Warning, Mary Jennerjohn, Mark Mcquilling, Ashley Purkey, Richard Scharnhorst, Mortaza ManiAbstract:Wind tunnels used to study Shock Wave boundary layer interactions (SBLIs) generally have small rectangular test sections that contain corner flow which affect the SBLIs. It is hypothesized that extending the width of the test section to increase the aspect ratio may deter corner flow from changing centerline flow. This study uses a Reynolds-averaged Navier-Stokes flow solver (Wind-US 3.0) with the Spalart-Allmaras turbulence model to investigate a Normal SBLI at Mach 1.6 in a 4.3 aspect ratio test section at two different Reynolds numbers: 1.62 · 10/ft for simulation A and 9.75 · 10/ft for simulation B. Analysis is focused on the spanwise composition of the SBLI, where results show the corner flows generate compression and reflected Shock Wave groups which disrupt spanwise uniformity. These Shocks sufficiently decrease the local Mach number upstream of the SBLI so that a lambda foot no longer exists away from the centerline. Away from the centerline the Shock structure changes as the main Shock migrates into the lower channel and a bow Shock forms at the Shock holder. This contradicts previous understanding of the lambda foot structure under a Normal Shock Wave in proximity to a corner flow.
Tsc Davidson - One of the best experts on this subject based on the ideXlab platform.
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influence of transition on the flow downstream of Normal Shock Wave boundary layer interactions
54th AIAA Aerospace Sciences Meeting, 2016Co-Authors: Tsc Davidson, H BabinskyAbstract:© 2016, Todd Davidson and Holger Babinsky. This paper reports recent work undertaken to investigate the interaction between Normal Shocks and laminar/transitional boundary layers. Laminar flow is desirable for its low skin friction, but problems are expected with Shock interactions as the boundary layer will be prone to separation. Experiments have been performed on a flat plate to explore the fundamental interaction without complex geometries and flow fields; and on a transonic aerofoil to research the influence of the imposed pressure field. The boundary layer state does not appear to significantly affect the downstream flow development, contrary to expectations, as the laminar separation is so small. On the aerofoil, a similar result is observed, with no change in the buffet onset Mach number and little change in the wake. This implies that Normal Shock Wave–laminar boundary layer interactions are not as detrimental as feared, and are in fact quite benign.
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transition location effects on Normal Shock Wave boundary layer interactions
53rd AIAA Aerospace Sciences Meeting, 2015Co-Authors: Tsc Davidson, H BabinskyAbstract:© 2015 by Todd Davidson and Holger Babinsky. Published by the American Institute of Aeronautics and Astronautics, Inc. This paper reports recent work undertaken to investigate the interaction between Normal Shocks and laminar/transitional boundary layers. Laminar ow is desirable for its low skin friction, but problems are expected with Shock interactions as the boundary layer will be prone to separation. Experiments have been performed on a at plate to explore the fundamental interaction without complex geometries and ow fields, and on a transonic aerofoil to research the in uence of the imposed pressure field. The boundary layer state before the Shock on the at plate does not appear to significantly affect the downstream ow development, contrary to expectations, as the laminar separation is so small. On the aerofoil, a similar result is observed, with no change in the buffet onset Mach number. This implies that Normal Shock Wave-laminar boundary layer interactions are not as detrimental as originally feared, and are in fact quite benign.
P R Ashill - One of the best experts on this subject based on the ideXlab platform.
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Normal Shock Wave turbulent boundary layer interactions in the presence of streamwise slots and grooves
Aeronautical Journal, 2002Co-Authors: A N Smith, H Babinsky, J L Fulker, P R AshillAbstract:The effect of streamwise slots and grooves on a Normal Shock Wave-turbulent boundary-layer interaction has been investigated experimentally at a Mach number of 1.3. The surface pressure distribution for the controlled interaction in the presence of slots featured a distinct plateau. This was due to a change in Shock structure from a typical unseparated Normal Shock Wave-boundary-layer interaction to a large bifurcated lambda type Shock pattern. Velocity measurements downstream of the slots revealed a strong spanwise variation of boundary-layer properties, whereas the modified Shock structure was found to be relatively two-dimensional. Cross flow measurements indicate that slots introduce streamwise vortices into the flow. When applied to an aerofoil, streamwise slots have the potential to reduce Wave drag while incurring only small viscous penalties. In the presence of grooves the interaction was initially found to be significantly different. A bifurcated Shock structure was observed but the trailing leg appeared stronger and featured a second lambda foot. Oil flow visualisation also revealed differences in the interactions, with the region of suction and blowing being limited to a smaller extent of the grooved control surface. The amount of crossflow present was reduced compared to the slotted control surface. By varying the internal geometry of the grooves it was found that the interaction could be modified to be similar to that in the presence of slots indicating that a more practical control device can be designed.
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control of Normal Shock Wave turbulent boundary layer interactions using streamwise grooves
40th AIAA Aerospace Sciences Meeting & Exhibit, 2001Co-Authors: A N Smith, H Babinsky, J L Fulker, P R AshillAbstract:The effect of streamwise slots on the interaction of a Normal Shock Wave / turbulent boundary layer has been investigated experimentally at a Mach number of 1.3. The surface pressure distribution for the controlled interaction was found to be significantly smeared, featuring a distinct plateau. This was due to a change in Shock structure from a typical unseparated Normal Shock Wave boundary layer interaction to a large bifurcated Lambda type Shock pattern. Boundary layer velocity measurements downstream of the slots revealed a strong spanwise variation of boundary layer properties whereas the modified Shock structure was relatively twodimensional. Oil flow visualisation indicated that in the presence of slots the boundary layer surface flow was highly three dimensional and confirmed that the effect of slots was mainly due to suction and blowing similar to that for passive control with uniform surface ventilation. Three hole probe measurements confirmed that the boundary layer was three dimensional and that the slots introduced vortical motion into the flowfield. Results indicate that when applied to an aerofoil, the control device has the potential to reduce Wave drag while incurring only small viscous penalties. The introduction of streamwise vorticity may also be beneficial to delay trailing edge separation and the device is thought to be capable of postponing buffet onset. © 2001 by A N Smith.
