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1
Sebastian, Jiss J., and Frank K. Lu. "Upstream-Influence Scaling of Fin-Induced Laminar Shockwave/Boundary-Layer Interactions." AIAA Journal 59, no. 5 (May 2021): 1861–64. http://dx.doi.org/10.2514/1.j059354.
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Sznajder, Janusz, and Tomasz Kwiatkowski. "EFFECTS OF TURBULENCE INDUCED BY MICRO VORTEX GENERATORS ON SHOCKWAVE – BOUNDARY LAYER INTERACTIONS." Journal of KONES. Powertrain and Transport 22, no. 2 (January 1, 2015): 241–48. http://dx.doi.org/10.5604/12314005.1165445.
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Gunasekaran, Humrutha, Thillaikumar Thangaraj, Tamal Jana, and Mrinal Kaushik. "Effects of Wall Ventilation on the Shock-wave/Viscous-Layer Interactions in a Mach 2.2 Intake." Processes 8, no. 2 (February 8, 2020): 208. http://dx.doi.org/10.3390/pr8020208.
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In order to achieve proficient combustion with the present technologies, the flow through an aircraft intake operating at supersonic and hypersonic Mach numbers must be decelerated to a low-subsonic level before entering the combustion chamber. High-speed intakes are generally designed to act as a flow compressor even in the absence of mechanical compressors. The reduction in flow velocity is essentially achieved by generating a series of oblique as well as normal shock waves in the external ramp region and also in the internal isolator region of the intake. Thus, these intakes are also referred to as mixed-compression intakes. Nevertheless, the benefits of shock-generated compression do not arise independently but with enormous losses because of the shockwave and boundary layer interactions (SBLIs). These interactions should be manipulated to minimize or alleviate the losses. In the present investigation a wall ventilation using a new cavity configuration (having a cross-section similar to a truncated rectangle with the top wall covered by a thin perforated surface is deployed underneath the cowl-shock impinging point of the Mach 2.2 mixed-compression intake. The intake is tested for four different contraction ratios of 1.16, 1.19, 1.22, and 1.25, with emphasis on the effect of porosity, which is varied at 10.6%, 15.7%, 18.8%, and 22.5%. The introduction of porosity on the surface covering the cavity has been proved to be beneficial in decreasing the wall static pressure substantially as compared to the plain intake. A maximum of approximately 24.2% in the reduction in pressure at the upstream proximal location of 0.48 L is achieved in the case of the wall-ventilated intake with 18.8% porosity, at the contraction ratio of 1.19. The Schlieren density field images confirm the efficacy of the 18.8% ventilation in stretching the shock trains and in decreasing the separation length. At the contraction ratios of 1.19, 1.22, and 1.25 (‘dual-mode’ contraction ratios), the controlled intakes with higher porosity reduce the pressure gradients across the shockwaves and thereby yields an ‘intake-start’ condition. However, for the uncontrolled intake, the ‘unstart’ condition emerges due to the formation of a normal shock at the cowl lip. Additionally, the cowl shock in the ‘unstart’ intake is shifted upstream because of higher downstream pressure.
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Zahrolayali, Nurfathin, Mohd Rashdan Saad, Azam Che Idris, and Mohd Rosdzimin Abdul Rahman. "Assessing the Performance of Hypersonic Inlets by Applying a Heat Source with the Throttling Effect." Aerospace 9, no. 8 (August 16, 2022): 449. http://dx.doi.org/10.3390/aerospace9080449.
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Utilization of a heat source to regulate the shock wave–boundary layer interaction (SWBLI) of hypersonic inlets during throttling was computationally investigated. A plug was installed at the intake isolator’s exit, which caused throttling. The location of the heat source was established by analysing the interaction of the shockwave from the compression ramp and the contact spot of the shockwave with that of the inlet cowl. Shockwave interaction inside the isolator was investigated using steady and transient cases. The present computational work was validated using previous experimental work. The flow distortion (FD) and total pressure recovery (TPR) of the inflows were also studied. We found that varying the size and power of the heat source influenced the shockwaves that originated around it and affected the SWBLI within the isolator. This influenced most of the performance measures. As a result, the TPR increased and the FD decreased when the heat source was applied. Thus, the use of a heat source for flow control was found to influence the performance of hypersonic intakes.
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Grilli, Muzio, Peter J. Schmid, Stefan Hickel, and Nikolaus A. Adams. "Analysis of unsteady behaviour in shockwave turbulent boundary layer interaction." Journal of Fluid Mechanics 700 (February 28, 2012): 16–28. http://dx.doi.org/10.1017/jfm.2012.37.
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AbstractThe unsteady behaviour in shockwave turbulent boundary layer interaction is investigated by analysing results from a large eddy simulation of a supersonic turbulent boundary layer over a compression–expansion ramp. The interaction leads to a very-low-frequency motion near the foot of the shock, with a characteristic frequency that is three orders of magnitude lower than the typical frequency of the incoming boundary layer. Wall pressure data are first analysed by means of Fourier analysis, highlighting the low-frequency phenomenon in the interaction region. Furthermore, the flow dynamics are analysed by a dynamic mode decomposition which shows the presence of a low-frequency mode associated with the pulsation of the separation bubble and accompanied by a forward–backward motion of the shock.
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Verma, S. B., and C. Manisankar. "Shockwave/Boundary-Layer Interaction Control on a Compression Ramp Using Steady Micro Jets." AIAA Journal 50, no. 12 (December 2012): 2753–64. http://dx.doi.org/10.2514/1.j051577.
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Prince, S. A., M. Vannahme, and J. L. Stollery. "Experiments on the hypersonic turbulent shock-wave/boundary-layer interaction and the effects of surface roughness." Aeronautical Journal 109, no. 1094 (April 2005): 177–84. http://dx.doi.org/10.1017/s0001924000000683.
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Abstract An experimental investigation was performed to study the effects of surface roughness on the Mach 8·2 hypersonic turbulent shockwave–boundary-layer interaction characteristics of a deflected control flap configuration. In particular, the surface pressure and heat transfer distribution along a quasi-2D ramp compression corner model was measured for flap angles between 0° and 38°, along with a Schlieren flow visualisation study. It was found that surface roughness, of scale 10% of the hinge-line boundary layer thickness, significantly increased the extent of the interaction, while increasing the magnitude of the peak pressure and heat flux just aft of reattachment. The incipient separation angle for a fully turbulent, Mach 8·2 boundary layer with a hinge line Reynolds number of 1·44 × 106, was estimated at 28-29°, reducing to between 19-22° with the introduction of laminar sub-layer scale surface roughness.
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Grisham, James R., Brian H. Dennis, and Frank K. Lu. "Incipient Separation in Laminar Ramp-Induced Shock-Wave/Boundary-Layer Interactions." AIAA Journal 56, no. 2 (February 2018): 524–31. http://dx.doi.org/10.2514/1.j056175.
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WU, MINWEI, and M. PINO MARTÍN. "Analysis of shock motion in shockwave and turbulent boundary layer interaction using direct numerical simulation data." Journal of Fluid Mechanics 594 (December 14, 2007): 71–83. http://dx.doi.org/10.1017/s0022112007009044.
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Direct numerical simulation data of a Mach 2.9, 24○ compression ramp configuration are used to analyse the shock motion. The motion can be observed from the animated DNS data available with the online version of the paper and from wall-pressure and mass-flux signals measured in the free stream. The characteristic low frequency is in the range of (0.007–0.013) U∞/δ, as found previously. The shock motion also exhibits high-frequency, of O(U∞/δ), small-amplitude spanwise wrinkling, which is mainly caused by the spanwise non-uniformity of turbulent structures in the incoming boundary layer. In studying the low-frequency streamwise oscillation, conditional statistics show that there is no significant difference in the properties of the incoming boundary layer when the shock location is upstream or downstream. The spanwise-mean separation point also undergoes a low-frequency motion and is found to be highly correlated with the shock motion. A small correlation is found between the low-momentum structures in the incoming boundary layer and the separation point. Correlations among the spanwise-mean separation point, reattachment point and the shock location indicate that the low-frequency shock unsteadiness is influenced by the downstream flow. Movies are available with the online version of the paper.
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Lee, S., and E. Loth. "Supersonic boundary-layer interactions with various micro-vortex generator geometries." Aeronautical Journal 113, no. 1149 (November 2009): 683–97. http://dx.doi.org/10.1017/s0001924000003353.
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Abstract Various types of micro-vortex generators (μVGs) are investigated for control of a supersonic turbulent boundary layer subject to an oblique shock impingement, which causes flow separation. The micro-vortex generators are embedded in the boundary layer to avoid excessive wave drag while still creating strong streamwise vortices to energise the boundary layer. Several different types of µVGs were considered including micro-ramps and micro-vanes. These were investigated computationally in a supersonic boundary layer at Mach 3 using monotone integrated large eddy simulations (MILES). The results showed that vortices generated from μVGs can partially eliminate shock induced flow separation and can continue to entrain high momentum flux for boundary-layer recovery downstream. The micro-ramps resulted in thinner downstream displacement thickness in comparison to the micro-vanes. However, the strength of the streamwise vorticity for the micro-ramps decayed faster due to dissipation especially after the shock interaction. In addition, the close spanwise distance between each vortex for the ramp geometry causes the vortex cores to move upwards from the wall due to induced upwash effects. Micro-vanes, on the other hand, yielded an increased spanwise spacing of the streamwise vortices at the point of formation. This resulted in streamwise vortices staying closer to the floor with less circulation decay, and the reduction in overall flow separation is attributed to these effects. Two hybrid concepts, named ‘thick-vane’ and ‘split-ramp’, were also studied where the former is a vane with side supports and the latter has a uniform spacing along the centreline of the baseline ramp. These geometries behaved similar to the micro-vanes in terms of the streamwise vorticity and the ability to reduce flow separation, but are more physically robust than the thin vanes.
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Yoon, B. K., M. K. Chung, and S. O. Park. "Comparisons between low Reynolds number two-equation models for computation of a shockwave-turbulent-boundary layer interaction." Aeronautical Journal 101, no. 1007 (September 1997): 335–45. http://dx.doi.org/10.1017/s0001924000066239.
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AbstractA comparative study is made on the performance of several low Reynolds number k-ε models and the k-ω model in predicting the shockwave-turbulent-boundary layer interaction over a supersonic compression ramp of 16°, 20° and 24° at a Mach numbers of 2.85, 2.79 and 2.84, respectively. The model equations are numerically solved by a higher order upwind scheme with the 3rd order MUSCL type TVD. The computational results reveal that all of the low Reynolds number k-ε models, particularly those employing y+ in their damping functions give erroneously large skin friction in the redeveloping region. It is also interesting to note that the k-ε models, when adjusted and based on DNS data, do not perform better, as expected, than the conventional low Reynolds number k-ε models. The k-ω model which does not adopt a low Reynolds number modification brings about reasonably accurate skin friction, but with a later onset of pressure rise. By recasting the ω equation into the general form of the ε equation, it is inferred that the turbulent cross diffusion term between k and ε is critical to guarantee better performance of the k-ω model for the skin friction prediction in the redeveloping region. Finally, an asymptotic analysis of a fully developed incompressible channel flow, with the k-ε and the k-ω models, reveals that the cross diffusion mechanism inherent in the k-ω model contributes to the better performance of the k-ω model.
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Pham, Harry T., Zachary N. Gianikos, and Venkateswaran Narayanaswamy. "Compression Ramp Induced Shock-Wave/Turbulent Boundary-Layer Interactions on a Compliant Material." AIAA Journal 56, no. 7 (July 2018): 2925–29. http://dx.doi.org/10.2514/1.j056652.
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KLASEBOER, EVERT, SIEW WAN FONG, CARY K. TURANGAN, BOO CHEONG KHOO, ANDREW J. SZERI, MICHAEL L. CALVISI, GEORGY N. SANKIN, and PEI ZHONG. "Interaction of lithotripter shockwaves with single inertial cavitation bubbles." Journal of Fluid Mechanics 593 (November 23, 2007): 33–56. http://dx.doi.org/10.1017/s002211200700852x.
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The dynamic interaction of a shockwave (modelled as a pressure pulse) with an initially spherically oscillating bubble is investigated. Upon the shockwave impact, the bubble deforms non-spherically and the flow field surrounding the bubble is determined with potential flow theory using the boundary-element method (BEM). The primary advantage of this method is its computational efficiency. The simulation process is repeated until the two opposite sides of the bubble surface collide with each other (i.e. the formation of a jet along the shockwave propagation direction). The collapse time of the bubble, its shape and the velocity of the jet are calculated. Moreover, the impact pressure is estimated based on water-hammer pressure theory. The Kelvin impulse, kinetic energy and bubble displacement (all at the moment of jet impact) are also determined. Overall, the simulated results compare favourably with experimental observations of lithotripter shockwave interaction with single bubbles (using laser-induced bubbles at various oscillation stages). The simulations confirm the experimental observation that the most intense collapse, with the highest jet velocity and impact pressure, occurs for bubbles with intermediate size during the contraction phase when the collapse time of the bubble is approximately equal to the compressive pulse duration of the shock wave. Under this condition, the maximum amount of energy of the incident shockwave is transferred to the collapsing bubble. Further, the effect of the bubble contents (ideal gas with different initial pressures) and the initial conditions of the bubble (initially oscillating vs. non-oscillating) on the dynamics of the shockwave-bubble interaction are discussed.
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Zhang, Zhi, Anna Hirahara, Yaohua Xue, and Naoki Uchiyama. "CFD simulation of shockwave-boundary layer interaction induced oscillation in NACA SC2-0714 transonic airfoil." Journal of Physics: Conference Series 2217, no. 1 (April 1, 2022): 012006. http://dx.doi.org/10.1088/1742-6596/2217/1/012006.
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Abstract To develop an advanced way of designing transonic airfoils for industry, a reliable and user-friendly mesh generator and a robust CFD solver are necessary. Especially, the prediction of unsteady shock buffet phenomenon is always a focus for the CFD simulation of transonic airfoils. In this study, BOXERmesh (an automatic mesh generator) and NEWT (a robust CFD solver), which are developed by CFS (Cambridge Flow Solutions) in collaboration with MHI (Mitsubishi Heavy Industries), are used to perform the CFD simulation of NACA SC2-0714 transonic airfoil. CFD simulation is conducted at two different attack angles with Mach number as 0.74 and Reynold number as 1.5×107 (non-buffet: α=2°; shock buffet: α=3°). A hexahedral dominant mesh is generated by BOXERmesh with the cell number as 5.52 million. Both unsteady RANS (URANS) and LES are performed using the CFD solver NEWT. Specifically, the governing equation is discretized by central differencing scheme with 2nd order accuracy in space by applying Swanson and Turkel type artificial viscosity, and the Adams-Bashford time integration with the dual-time stepping method is applied for temporal discretization in the density-based solver. Results show CFD simulation could reproduce the time averaged chordwise distribution of pressure coefficient at the two conditions. Both URANS and LES successfully capture the unsteady shock buffet phenomenon when increasing attack angle from 2° to 3°. However, the calculated peak oscillation location and the shock buffet frequency are different between URANS and LES. Applying the same mesh resolution, URANS performances better than LES, with the deviation of the shock buffet frequency less than 6% (Exp.: 69 Hz; URANS: 73 Hz). The reason is considered as the wall-normal mesh refinement near the airfoil surface (y+∼66) being not enough for LES to accurately resolve the turbulence scale and capture the boundary layer separation behavior. On the other hand, URANS is thought to be enough to reproduce the periodic moving of the onset of boundary separation and to predict the main characteristics of the shock buffet phenomenon.
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Wang, Lican, Yilong Zhao, Qiancheng Wang, Yuxin Zhao, Ruoling Zhang, and Li Ma. "Three-dimensional characteristics of crossing shock wave/turbulent boundary layer interaction in a double fin with and without micro-ramp control." AIP Advances 12, no. 9 (September 1, 2022): 095309. http://dx.doi.org/10.1063/5.0102986.
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The three-dimensional (3D) interactions between crossing shock waves and a turbulent boundary layer (CSWBLI) inside a symmetric double fin are experimentally studied using nanoparticle-based planar laser scattering, supersonic particle image velocimetry, and surface oil visualization. The possibility of controlling the separated flow generated by CSWBLI is considered by employing micro-ramp vortex generators. First, the fractal dimension, velocity profile, and logarithmic law of the incoming turbulent boundary layer at Mach number 2.8 are examined. Then, the flow structure and velocity distribution, which have seldom been presented in previous experiments, are measured in high resolution. The 3D behavior of the boundary layer after CSWBLI shows that the boundary layer becomes thicker behind the shock wave and converges toward the symmetry plane of the double fin. The converged effect contributes to the largest thickness of the boundary layer in the symmetry plane accompanied with a separation region near the wall. Introduction of seven equidistant micro-ramps upstream of the double fin is proved to suppress the separation region, where the arc-like vortices generated by the middle micro-ramps are found to be more sustainable along the streamwise direction. The micro-ramps can increase the momentum exchange between the boundary layer and the surrounding mainstream. At the same time, the momentum exchange induced by the micro-ramps decreases the flow velocity outside the converged region in comparison with the configuration without micro-ramps. The results obtained in this paper can provide an experimental insight into the 3D physical phenomena existing in the CSWBLI and its flow control.
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Xiang, Gaoming, Daiwei Li, Junqin Chen, Arpit Mishra, Georgy Sankin, Xuning Zhao, Yuqi Tang, Kevin Wang, Junjie Yao, and Pei Zhong. "Dissimilar cavitation dynamics and damage patterns produced by parallel fiber alignment to the stone surface in holmium:yttrium aluminum garnet laser lithotripsy." Physics of Fluids 35, no. 3 (March 2023): 033303. http://dx.doi.org/10.1063/5.0139741.
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Recent studies indicate that cavitation may play a vital role in laser lithotripsy. However, the underlying bubble dynamics and associated damage mechanisms are largely unknown. In this study, we use ultra-high-speed shadowgraph imaging, hydrophone measurements, three-dimensional passive cavitation mapping (3D-PCM), and phantom test to investigate the transient dynamics of vapor bubbles induced by a holmium:yttrium aluminum garnet laser and their correlation with solid damage. We vary the standoff distance ( SD) between the fiber tip and solid boundary under parallel fiber alignment and observe several distinctive features in bubble dynamics. First, long pulsed laser irradiation and solid boundary interaction create an elongated “pear-shaped” bubble that collapses asymmetrically and forms multiple jets in sequence. Second, unlike nanosecond laser-induced cavitation bubbles, jet impact on solid boundary generates negligible pressure transients and causes no direct damage. A non-circular toroidal bubble forms, particularly following the primary and secondary bubble collapses at SD = 1.0 and 3.0 mm, respectively. We observe three intensified bubble collapses with strong shock wave emissions: the intensified bubble collapse by shock wave, the ensuing reflected shock wave from the solid boundary, and self-intensified collapse of an inverted “triangle-shaped” or “horseshoe-shaped” bubble. Third, high-speed shadowgraph imaging and 3D-PCM confirm that the shock origins from the distinctive bubble collapse form either two discrete spots or a “smiling-face” shape. The spatial collapse pattern is consistent with the similar BegoStone surface damage, suggesting that the shockwave emissions during the intensified asymmetric collapse of the pear-shaped bubble are decisive for the solid damage.
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Nurfathin Zahrolayali, Mohd Rashdan Saad, Azam Che Idris, and Mohd Rosdzimin Abdul Rahman. "Numerical Investigation of the Hypersonic Inlet under Throttling with Heat Source." CFD Letters 14, no. 10 (October 28, 2022): 79–86. http://dx.doi.org/10.37934/cfdl.14.10.7986.
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The influence of using a heat source to manage the shock wave boundary layer interactions (SWBLI) at the hypersonic inlet under throttling were studied numerically. This hypersonic inlet was created for a fluid flow Mach number of 5. The throttling was induced by a plug placed near the intake isolator's outlet. The study's parameters included the heat source power and size. The intake performance indicators were the total pressure recovery and the flow distortion. The position of the heat source was determined by studying the interplay of the shock waves from the compression ramp. The results demonstrated the existence of the shock waves at the heat source, and its influences on the SWBLI inside the isolator. This behaviour, led to an increase in the total pressure recovery and reduction of the flow distortion.
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Zuo, Feng-Yuan. "Hypersonic Shock Wave/Turbulent Boundary Layer Interaction over a Compression Ramp." AIAA Journal, January 2, 2023, 1–17. http://dx.doi.org/10.2514/1.j062521.
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A parametric study of ramp-induced planar shock-wave/turbulent-boundary-layer interactions (SBLIs) is carried out at hypersonic conditions (Mach number 6.0) by means of numerical simulation of the Reynolds-averaged Navier–Stokes (RANS) equations, with the eventual goal of establishing wall temperature and Reynolds number effects. Comparison with available experimental data shows that RANS is capable of predicting the main features of hypersonic oblique SBLI, namely, typical size and distribution of the wall-surface pressure, and heat transfer. A large number of flow cases, at low ([Formula: see text]) and high Reynolds number ([Formula: see text]), were computed to examine the scaling of the heat transfer over a wide range of wall temperatures. As expected, the interaction zone of hypersonic ramp-induced SBLI is reduced as the wall is cooled. A simple power law for heat transfer originally introduced by Back and Cuffel (AIAA Journal, Vol. 8, No. 10, 1970, pp. 1871–1873) is here considered to account for hypersonic ramp-induced SBLI, which is found to successfully collapse the data to the distributions obtained for supersonic, cold/hot interactions.
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Zhou, Jin, Yu Liu, and Zhi-yong Lin. "Numerical study on the standing morphology of an oblique detonation wave under the influence of an incoming boundary layer." Open Physics 13, no. 1 (January 1, 2015). http://dx.doi.org/10.1515/phys-2015-0007.
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AbstractThe influence of an incoming boundary layer to the standing morphology of an oblique detonation wave (ODW) induced by a compression ramp is numerically studied in this paper. The Spalart-Allmaras (SA) turbulence model is used to perform simulation of detonationboundary- layer interactions. Three different wall conditions are applied to realize control on the boundary-layer separation scales. Accordingly, different standing morphologies of the ODWs are obtained, including smooth ODW (without transverse wave) under no-slip, adiabatic wall condition with large-scale separation, abrupt ODW (with transverse wave) under no-slip, cold wall condition with moderate-scale separation, and bow-shaped detached ODW under slipwall condition without a boundary layer.
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Pandey, Anshuman, Katya M. Casper, Steven J. Beresh, Rajkumar Bhakta, and Russell Spillers. "Hypersonic Fluid–Structure Interaction on a Cone–Slice–Ramp Geometry." AIAA Journal, February 12, 2023, 1–17. http://dx.doi.org/10.2514/1.j062733.
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Fluid–structure interactions were measured between a representative control surface and the hypersonic flow deflected by it. The control surface is simplified as a spanwise finite ramp placed on a longitudinal slice of a cone. The front surface of the ramp contains a thin panel designed to respond to the unsteady fluid loading arising from the shock-wave/boundary-layer interactions. Experiments were conducted at Mach 5 and Mach 8 with ramps of different angles. High-speed schlieren captured the unsteady flow dynamics and accelerometers behind the thin panel measured its structural response. Panel vibrations were dominated by natural modes that were excited by the broadband aerodynamic fluctuations arising in the flowfield. However, increased structural response was observed in two distinct flow regimes: 1) attached or small separation interactions, where the transitional regime induced the strongest panel fluctuations. This was in agreement with the observation of increased convective undulations or bulges in the separation shock generated by the passage of turbulent spots, and 2) large separated interactions, where shear layer flapping in the laminar regime produced strong panel response at the flapping frequency. In addition, panel heating during the experiment caused a downward shift in its natural mode frequencies.
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D’Aguanno, A., P. Quesada Allerhand, F. F. J. Schrijer, and B. W. van Oudheusden. "Characterization of shock-induced panel flutter with simultaneous use of DIC and PIV." Experiments in Fluids 64, no. 1 (January 2023). http://dx.doi.org/10.1007/s00348-022-03551-1.
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AbstractIn this experimental study, panel flutter induced by an impinging oblique shockwave is investigated at a freestream Mach number of 2, using the combination of planar particle image velocimetry (PIV) and stereographic digital image correlation (DIC) to obtain simultaneous full-field structural displacement and flow velocity measurements. High-speed cameras are employed to obtain a time-resolved description of the panel motion and the shockwave-boundary layer interaction (SWBLI). In order to prevent interference between the PIV and DIC systems, an optical isolation is implemented using fluorescent paint, dedicated light sources, and camera lens filters. The effect of the panel motion on the SWBLI behavior is assessed, by comparing it with the SWBLI on a rigid wall. The results show that panel oscillations occur with a maximum amplitude of ten times the panel thickness. The dominant frequencies observed in the panel oscillation (424 Hz and 1354 Hz) match the main spectral content of the reflected shockwave position. A further POD analysis of the panel displacement spatial distribution shows that these two frequency contributions are well captured by the first two POD modes, which correspond, respectively, to a first and a third bending mode shape and account for 92% of the total oscillation energy. The fluid-structure coupling is studied by identifying, in the flow, the regions of maximum correlation between the panel displacement and the flow velocity fluctuations. The results obtained prove that the inviscid flow region upstream of the SWBLI is perfectly in phase with the panel oscillation, while the downstream region has a delay of one quarter of the flutter cycle.
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Ramaswamy, Deepak Prem, and Anne-Marie Schreyer. "Separation control with elliptical air-jet vortex generators." Experiments in Fluids 64, no. 5 (May 2023). http://dx.doi.org/10.1007/s00348-023-03637-4.
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AbstractThe flow organisation of air-jet vortex generators (AJVGs) of elliptical cross sections and their control effectiveness on a $$24^{\circ }$$ 24 ∘ -compression-ramp-induced shock-wave/boundary-layer interaction was analysed on the basis of experiments at $$M_{\infty } = 2.52$$ M ∞ = 2.52 and $$Re_{\theta _c} = 8225$$ R e θ c = 8225 . We investigated a circular orifice and two elliptical orifices of aspect ratios 0.5 and 2; all characterised by the same hydraulic diameter. Measurements of separation lengths from oil-flow visualisation and PIV reveal that elliptical AJVGs achieve a $$25\%$$ 25 % reduction in total separation length, which constitutes a strong improvement over the $$17\%$$ 17 % reduction achieved with the commonly used circular AJVGs. The jet-induced structures from elliptical AJVGs penetrate on average $$25\%$$ 25 % farther into the boundary layer. However, the lateral spread is limited to a maximum value equal to the inter-jet spacing in the control array, which highlights the onset of jet/jet interactions between adjacent jets in the array. A consequence of these interactions is better flow entrainment for the elliptical cases, as observed in the mean boundary-layer velocity profiles and an improved turbulent mixing (indicated by an increase in Reynolds-shear-stress magnitude).
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Das, Dipankar, Siddesh Desai, and Vinayak Kulkarni. "Comparative studies of shock-wave boundary-layer interactions in Earth and Mars atmospheres." Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, April 22, 2021, 095441002110118. http://dx.doi.org/10.1177/09544100211011829.
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Investigations of ramp-induced shock-wave boundary-layer interaction have been carried out for real gas flows of air and carbon dioxide through hypersonic laminar flow simulations corresponding to Earth and Mars atmospheres. An in-house-developed solver, which accounts for the real gas effects, has been employed for these studies. Effects of various parameters like wall temperature, freestream stagnation enthalpy, freestream Mach number, and blunt leading edge are explored on the intensity of shock-wave boundary-layer interaction (SWBLI). In either case, an increase in separation length is observed with an increase in wall temperature and a decrease in Mach number as well as freestream stagnation enthalpy. Here, the intensity of alteration is always noted to have a higher percentage for the Mars gas model. Further, separation length is found to be almost equal for the same wall to total temperature ratio in both of the flow mediums. The present study also affirms the fact that the leading edge bluntness can be used as a tool to reduce the size of the separation region in these planetary atmospheres. Revised correlations have been proposed for hypersonic Earth atmospheric flow with real gas effects to predict the extent of upstream influence and separation bubble size. The outcomes of simulations have also helped to device new correlations for these flow features of SWBLI for Mars atmospheric conditions. In all, the need for consideration of real gas effects and an exclusive real gas flow solver for the Mars atmosphere are the prominent recommendations of current studies.
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Zangeneh, Rozie. "Development of a New Algorithm for Modeling Viscous Transonic Flow on Unstructured Grids at High Reynolds Numbers." Journal of Fluids Engineering 143, no. 2 (October 26, 2020). http://dx.doi.org/10.1115/1.4048611.
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Abstract This study investigates a new algorithm for modeling viscous transonic flow at high Reynolds number cases suitable for unstructured grids. The challenge of modeling viscous transonic flow around airfoils becomes intense at high Reynolds number cases due to a variety of flow regimes encountered, such as boundary layer growth and the shockwave/turbulent boundary-layer interaction, accompanied by large separation bubble. Therefore, it is highly demanded to develop robust and efficient models that can capture the shock-induced problems of turbulent flows for aircraft design purposes. The new model is essentially a hybrid algorithm to address the conflict between turbulence modeling and shock-capturing requirements. A skew-symmetric form of a collocated finite volume scheme with minimum aliasing errors was implemented to model the turbulent region in the combination of a semidiscrete, central difference scheme to capture discontinuities with adequately low numerical dissipation for the minimal effect on turbulent flows. To evaluate the effectiveness of the model, it was tested in three conventional cases. The computational results are close to measured data for predicting the shock locations. This implies that the model is able to predict the scale of the separation bubble and the main characteristics of turbulent transonic flow adequately.
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Hu, Weibo, Stefan Hickel, and Bas W. van Oudheusden. "Unsteady mechanisms in shock wave and boundary layer interactions over a forward-facing step." Journal of Fluid Mechanics 949 (September 21, 2022). http://dx.doi.org/10.1017/jfm.2022.737.
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The flow over a forward-facing step (FFS) at $Ma_\infty =1.7$ and $Re_{\delta _0}=1.3718\times 10^{4}$ is investigated by well-resolved large-eddy simulation. To investigate effects of upstream flow structures and turbulence on the low-frequency dynamics of the shock wave/boundary layer interaction (SWBLI), two cases are considered: one with a laminar inflow and one with a turbulent inflow. The laminar inflow case shows signs of a rapid transition to turbulence upstream of the step, as inferred from the streamwise variation of $\langle C_f \rangle$ and the evolution of the coherent vortical structures. Nevertheless, the separation length is more than twice as large for the laminar inflow case, and the coalescence of compression waves into a separation shock is observed only for the fully turbulent inflow case. The dynamics at low and medium frequencies is characterized by a spectral analysis, where the lower frequency range is related to the unsteady separation region, and the intermediate one is associated with the shedding of shear layer vortices. For the turbulent inflow case, we furthermore use a three-dimensional dynamic mode decomposition to analyse the individual contributions of selected modes to the unsteadiness of the SWBLI. The separation shock and Görtler-like vortices, which are induced by the centrifugal forces in the separation region, are strongly correlated with the low-frequency unsteadiness in the current FFS case. Similarly as observed previously for the backward-facing steps, we observe a slightly higher non-dimensional frequency (based on the separation length) of the low-frequency mode than for SWBLI in flat plate and ramp configurations.
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Sebastian, R., and A. M. Schreyer. "Influence of jet spacing in spanwise-inclined jet injection in supersonic crossflow." Journal of Fluid Mechanics 946 (August 11, 2022). http://dx.doi.org/10.1017/jfm.2022.597.
Full textAbstract:
In separation control with air-jet vortex generators in supersonic flow, the spacing of the jets in a row array has a crucial effect on the control effectiveness. Previous experimental studies have revealed the overall influence of jet spacing on air-jet vortex generator controlled shock-induced separation zones, focusing mostly on the separation regions. There is, however, a gap in knowledge regarding the mechanisms leading to these changes in control effectiveness, particularly on the interaction between multiple jets and the details of downstream flow evolution. Therefore, the objective of the current study is to provide detailed information on the underlying flow dynamics associated with the injection of a row of spanwise-inclined jets into a supersonic turbulent boundary layer – and in particular the effects of different jet spacings. Four different spacings were studied with large-eddy simulations. In addition, we performed oil-flow and schlieren visualizations of separation control in a 24 $^\circ$ compression–ramp interaction with different jet-spacing configurations to validate and discuss our conclusions regarding the effects of jet/jet interactions on the separation-control effectiveness. We analyse the influence of jet spacing on the flow topology, induced vortical structures, and boundary-layer statistics. The jet-induced major vortex pair is the dominant flow structure energizing the near-wall boundary layer. The paper details how interactions amongst adjacent vortex pairs differ for varying jet spacings, thus influencing the momentum transfer and eventually the control efficiency. The dynamic behaviour of the flow was analysed using a three-dimensional dynamic mode decomposition. The resulting insights are key to the development of efficient control set-ups.
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Ramaswamy, Deepak Prem, and Anne-Marie Schreyer. "Effects of Jet-to-Jet Spacing of Air-Jet Vortex Generators in Shock-Induced Flow-Separation Control." Flow, Turbulence and Combustion, April 12, 2022. http://dx.doi.org/10.1007/s10494-022-00324-y.
Full textAbstract:
AbstractExperiments were carried out to assess the influence of spanwise spacing between adjacent orifices of an air-jet vortex-generator (AJVG) array on their separation-control effectiveness. The array was applied to a 24° compression-ramp-induced shock-wave / turbulent boundary-layer interaction at $$M_{\infty } = 2.52$$ M ∞ = 2.52 and $$Re_{\theta } = 8225$$ R e θ = 8225 . Three spanwise oriented AJVG arrays of small, intermediate, and large jet spacings were studied. Their influence on the mean-flow organisation and turbulence quantities was assessed using flow visualisations and planar particle image velocimetry across multiple measurement planes. The streamwise vortices induced by the AJVGs incited different control effects depending on the degree of interaction between adjacent vortices. The array with intermediate spacing achieved the most favourable effects with reductions in separation length and area of about $$25\%$$ 25 % and $$52\%$$ 52 % , respectively. This reduction was brought about by the formation of stable, interacting streamwise-elongated coherent vortices downstream of jet injection and the associated entrainment of high-momentum fluid. The smallest jet spacing incites vortex interactions to adverse strength, breaking up the coherent structures and even increasing separation. The AJVGs with the largest spacing display characteristics similar to single jets in crossflow, with only a local modification of the separation region. Turbulent quantities are amplified both by the jets and the separation-inducing shock; AJVG control reduces the amplification across the shock-wave/boundary-layer interaction and the intermediate jet spacing is most effective also here.