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Статті в журналах з теми "Shockwave Boundary Layer Interactions"

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Автор: Grafiati

Опубліковано: 6 вересня 2023

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1

Chokani, N., and L. C. Squire. "Transonic shockwave/turbulent boundary layer interactions on a porous surface." Aeronautical Journal 97, no. 965 (May 1993): 163–70. http://dx.doi.org/10.1017/s0001924000026117.

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AbstractTransonic shockwave/turbulent boundary layer interactions on a porous surface above a closed plenum chamber have been studied experimentally in the choked flow of a windtunnel test-section. The equivalent freestream Mach number is 0.76 and results were obtained for three shock strengths. Without the porous surface the Mach numbers ahead of the shock were 1.13, 1.18 and 1.26. The respective shock Mach numbers with the porous surface were 1.10, 1.11 and 1.19. Laser holographic interferometry results are used to measure the density flowfield and examine the nature of the interaction. These results show that the interaction on the porous surface is modified by a thin shear layer adjacent to the surface and the weakening of the Shockwave is attributed to this. The interaction was also studied by solving the two-dimensional Reynolds-averaged Navier-Stokes equations together with the two-layer algebraic eddy-viscosity model of Baldwin-Lomax modified with appropriate corrections for surface transpiration. The computed results show excellent agreement with the experimental data. The examination of these numerical results shows that the surface transpiration occurs at a low subsonic velocity and suggests that the effect of the transpiration through the porous surface on the interaction may be optimised.

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2

Chand, S. V. S. A. Hema Sai. "Transonic shockwave/boundary layer interactions on NACA 5 series -24112." International Journal of Current Engineering and Technology 2, no. 2 (January 1, 2010): 629–34. http://dx.doi.org/10.14741/ijcet/spl.2.2014.120.

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3

Hanna, Rebecca L. "Hypersonic shockwave/turbulent boundary-layer interactions on a porous surface." AIAA Journal 33, no. 10 (October 1995): 1977–79. http://dx.doi.org/10.2514/3.12755.

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4

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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5

Délery, J. M. "Shock phenomena in high speed aerodynamics: still a source of major concern." Aeronautical Journal 103, no. 1019 (January 1999): 19–34. http://dx.doi.org/10.1017/s0001924000065076.

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Анотація:

Abstract Shockwaves are present in a flow as soon as the Mach number becomes supersonic. Being viscous phenomena, Shockwaves are a source of drag which can be predominant when the Mach number is significantly higher than one. In supersonic air intakes, the production of entropy by shocks is felt as a loss in efficiency. At high Mach numbers, Shockwaves produce a considerable temperature rise leading to severe heating problems, complicated by real gas effects. The intersection - or interference - of two shocks gives rise to complex wave patterns containing slip-lines and associated shear layers whose impingement on a nearby surface can cause detrimental pressure and heat transfer loads. The impact of a Shockwave on a boundary layer is the origin of strong viscous interactions which remain a limiting factor in the design of transonic wings, supersonic air intakes, propulsive nozzles and compressor cascades. More effort is needed to improve prediction of these interactions and to devise new techniques to control such phenomena.

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6

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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7

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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8

Hamed, A., and J. S. Shang. "Survey of validation data base for shockwave boundary-layer interactions in supersonic inlets." Journal of Propulsion and Power 7, no. 4 (July 1991): 617–25. http://dx.doi.org/10.2514/3.23370.

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9

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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10

Kalra, Chiranjeev S., Sohail H. Zaidi, Richard B. Miles, and Sergey O. Macheret. "Shockwave–turbulent boundary layer interaction control using magnetically driven surface discharges." Experiments in Fluids 50, no. 3 (August 18, 2010): 547–59. http://dx.doi.org/10.1007/s00348-010-0898-9.

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11

Piovesan, Tommaso, Andrea Magrini, and Ernesto Benini. "Accurate 2-D Modelling of Transonic Compressor Cascade Aerodynamics." Aerospace 6, no. 5 (May 19, 2019): 57. http://dx.doi.org/10.3390/aerospace6050057.

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Modern aeronautic fans are characterised by a transonic flow regime near the blade tip. Transonic cascades enable higher pressure ratios by a complex system of shockwaves arising across the blade passage, which has to be correctly reproduced in order to predict the performance and the operative range. In this paper, we present an accurate two-dimensional numerical modelling of the ARL-SL19 transonic compressor cascade. A large series of data from experimental tests in supersonic wind tunnel facilities has been used to validate a computational fluid dynamic model, in which the choice of turbulence closure resulted critical for an accurate reproduction of shockwave-boundary layer interaction. The model has been subsequently employed to carry out a parametric study in order to assess the influence of main flow variables (inlet Mach number, static pressure ratio) and geometric parameters (solidity) on the shockwave pattern and exit status. The main objectives of the present work are to perform a parametric study for investigating the effects of the abovementioned variables on the cascade performance, in terms of total-pressure loss coefficient, and on the shockwave pattern and to provide a quite large series of data useful for a preliminary design of a transonic compressor rotor section. After deriving the relation between inlet and exit quantities, peculiar to transonic compressors, exit Mach number, mean exit flow angle and total-pressure loss coefficient have been examined for a variety of boundary conditions and parametrically linked to inlet variables. Flow visualisation has been used to describe the shock-wave pattern as a function of the static pressure ratio. Finally, the influence of cascade solidity has been examined, showing a potential reduction of total-pressure loss coefficient by employing a higher solidity, due to a significant modification of shockwave system across the cascade.

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12

Sun, Quan, Wei Cui, Ying-Hong Li, Bang-Qin Cheng, Di Jin, and Jun Li. "Shockwave—boundary layer interaction control by plasma aerodynamic actuation: An experimental investigation." Chinese Physics B 23, no. 7 (July 2014): 075210. http://dx.doi.org/10.1088/1674-1056/23/7/075210.

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13

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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14

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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15

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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16

Sznajder, Janusz, and Tomas Kwiatkowski. "ANALYSIS OF EFFECTS OF SHAPE AND LOCATION OF MICRO-TURBULATORS ON UNSTEADY SHOCKWAVE-BOUNDARY LAYER INTERACTIONS IN TRANSONIC FLOW." Journal of KONES. Powertrain and Transport 23, no. 2 (January 1, 2016): 373–80. http://dx.doi.org/10.5604/12314005.1213755.

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17

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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18

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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19

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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20

Deng, Feng, Shenghua Zhang, and Ning Qin. "Closed-Loop Control of Transonic Buffet Using Active Shock Control Bump." Aerospace 10, no. 6 (June 4, 2023): 537. http://dx.doi.org/10.3390/aerospace10060537.

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At transonic flight conditions, the buffet caused by the shockwave/boundary-layer interaction can degrade aircraft performance and even threaten their safety. In this paper, a closed-loop control using an active shock control bump (SCB) has been proposed to suppress the buffet on a supercritical airfoil flying at transonic speeds. A closed-loop control law is designed by using the lift coefficient as the feedback signal and using the bump height as the control variable. The unsteady numerical simulations show that the buffet can be effectively suppressed by an optimal combination of the parameters of the control law, namely the gain and the delay time. Furthermore, the buffet control effectiveness is still acceptably constrained by a prescribed maximum bump height, which is believed to be practically important. In addition to being able to achieve both wave drag reduction and buffet alleviation, the active SCB is less sensitive to the parameters of the control law and has a shorter response time in comparison with the reference active trailing edge flap.

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21

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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22

Cutler, A. D., and P. Bradshaw. "Strong vortex/boundary layer interactions." Experiments in Fluids 14, no. 5 (April 1993): 321–32. http://dx.doi.org/10.1007/bf00189490.

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23

Cutler, A. D., and P. Bradshaw. "Strong vortex/boundary layer interactions." Experiments in Fluids 14, no. 6 (May 1993): 393–401. http://dx.doi.org/10.1007/bf00190193.

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24

Wasistho, B. "Transpiration Induced Shock Boundary-Layer Interactions." Journal of Fluids Engineering 128, no. 5 (January 28, 2006): 976–86. http://dx.doi.org/10.1115/1.2236127.

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Steady and unsteady shock boundary-layer interactions are studied numerically by solving the two-dimensional time-dependent Navier-Stokes equations. To validate the numerical method, the steady interaction is compared with measurements and other numerical results reported in the literature. The numerical study of the steady interaction leads to a suitable method for transpiration boundary conditions. The method applies to unsteady flows as well. Using the validated numerical method, we show that an unsteady shock boundary-layer interaction can occur in a supersonic flow over a flat plate subjected to suction and blowing from the opposite side of the plate, even though the imposed transpiration is steady. Depending on the Mach number, the Reynolds number, the distance of the transpiration boundary to the lower wall, and the transpiration profile, the unsteadiness can be inviscid or viscous dominated. The viscous effect is characterized by the occurrence of self-excited vortex shedding. A criterion for the onset of vortex shedding for internal compressible flows is also proposed.

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25

Liu, Jing Yuan, and Chun Hian Lee. "Development of A Two-Equation Turbulence Model for Hypersonic Shock Wave and Turbulent Boundary Layer Interaction." Applied Mechanics and Materials 66-68 (July 2011): 1868–73. http://dx.doi.org/10.4028/www.scientific.net/amm.66-68.1868.

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For hypersonic compressible turbulence, the correlations with respect to the density fluctuation must not be neglected. A Reynolds averaged K-ε model is proposed in the present paper to include these correlations, together with the Reynolds averaged Navier-Stokes equations to describe the mean flowfield. The K-equation is obtained from Reynolds averaged single-point second moment equations which are deduced from the instantaneous compressible Navier-Stokes equations. Under certain hypotheses and scales estimation of the compressible terms, the K-equation is simplified. The correlation terms of the fluctuation field appearing in the resulting K-equation, together with a conventional form of the ε-equation, are thus correlated with the variables in the average field. The new modeling coefficients of closure terms are optimized by computing the hypersonic turbulent flat-plate measured by Coleman and Stollery [J. Fliud Mech., Vol. 56 (1972), p. 741]. The proposed model is then applied to simulate hypersonic turbulent flows over a wedge compression corner angle of 34 degree. The predicting results compare favorably with the experimental results. Also, comparisons are made with other turbulence models. Additionally, an entropy modification function of Harten-Yee’s TVD scheme is introduced to reduce artificial diffusion near boundary layers and provide the required artificial diffusion to capture the shockwaves simultaneously.

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26

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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27

Li, Guangning, Dinesh Bhatia, Min Xu, and Jian Wang. "Grid Convergence Analysis for MUSCL-based Numerical Scheme in Shockwave-containing Flows." MATEC Web of Conferences 257 (2019): 02001. http://dx.doi.org/10.1051/matecconf/201925702001.

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This paper investigated the influence of limiter functions widely utilized in MUSCL-type (Monotone Upstream-centred Schemes for Conservation Laws) upwind numerical schemes on the solution accuracy of shockwave-containing flows. An incident shock interacting with laminar boundary layer developed on a flat plate was numerically simulated with the in-house developed code. A mixed-order grid convergence study was performed to assess the spatial errors of different limiters in simulating the selected shockwave-containing flow on flat plate. The conclusions are that, limiter functions implemented in the current in-house code play the critical roles in accurately predicting shockwave-containing flows. The mixed-order error estimator based on grid convergence study was proved to be applicable to evaluate the spatial errors of shockwave-containing flows, where the shock could reduce the nominal second- or third-order accuracy to first-order. The mixed-order estimator is conservative in the sense that the actual error is less than the error estimated, in the examined case.

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28

Zhang, Chuanhong, and Zhiwei Shi. "Nonlinear interactions in a hypersonic boundary layer." AIP Advances 11, no. 3 (March 1, 2021): 035104. http://dx.doi.org/10.1063/5.0044143.

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29

Mowatt, S., and B. Skews. "Three dimensional shock wave/boundary layer interactions." Shock Waves 21, no. 5 (May 8, 2011): 467–82. http://dx.doi.org/10.1007/s00193-011-0322-2.

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30

Gaitonde, Datta V. "Progress in shock wave/boundary layer interactions." Progress in Aerospace Sciences 72 (January 2015): 80–99. http://dx.doi.org/10.1016/j.paerosci.2014.09.002.

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31

Howes, F. A. "Some models of shock-boundary layer interactions." Journal of Mathematical Analysis and Applications 138, no. 1 (February 1989): 199–208. http://dx.doi.org/10.1016/0022-247x(89)90330-2.

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32

Kluwick, Alfred, and Marcus Wrabel. "Shock boundary layer interactions in dense gases." PAMM 4, no. 1 (December 2004): 444–45. http://dx.doi.org/10.1002/pamm.200410203.

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33

Carpenter, D. L., and J. Lemaire. "The Plasmasphere Boundary Layer." Annales Geophysicae 22, no. 12 (December 22, 2004): 4291–98. http://dx.doi.org/10.5194/angeo-22-4291-2004.

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Abstract. As an inner magnetospheric phenomenon the plasmapause region is of interest for a number of reasons, one being the occurrence there of geophysically important interactions between the plasmas of the hot plasma sheet and of the cool plasmasphere. There is a need for a conceptual framework within which to examine and discuss these interactions and their consequences, and we therefore suggest that the plasmapause region be called the Plasmasphere Boundary Layer, or PBL. Such a term has been slow to emerge because of the complexity and variability of the plasma populations that can exist near the plasmapause and because of the variety of criteria used to identify the plasmapause in experimental data. Furthermore, and quite importantly in our view, a substantial obstacle to the consideration of the plasmapause region as a boundary layer has been the longstanding tendency of textbooks on space physics to limit introductory material on the plasmapause phenomenon to zeroth order descriptions in terms of ideal MHD theory, thus implying that the plasmasphere is relatively well understood. A textbook may introduce the concept of shielding of the inner magnetosphere from perturbing convection electric fields, but attention is not usually paid to the variety of physical processes reported to occur in the PBL, such as heating, instabilities, and fast longitudinal flows, processes which must play roles in plasmasphere dynamics in concert with the flow regimes associated with the major dynamo sources of electric fields. We believe that through the use of the PBL concept in future textbook discussions of the plasmasphere and in scientific communications, much progress can be made on longstanding questions about the physics involved in the formation of the plasmapause and in the cycles of erosion and recovery of the plasmasphere. Key words. Magnetospheric physics (plasmasphere; plasma convection; MHD waves and instabilities)

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34

KHOO, BOO CHEONG, DEEPAK ADIKHARI, SIEW WAN FONG, and EVERT KLASEBOER. "MULTIPLE SPARK-GENERATED BUBBLE INTERACTIONS." Modern Physics Letters B 23, no. 03 (January 30, 2009): 229–32. http://dx.doi.org/10.1142/s0217984909018072.

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The complex interactions of two and three spark-generated bubbles are studied using high speed photography. The corresponding simulations are performed using a 3D Boundary Element Method (BEM) code. The bubbles generated are between 3 to 5 mm in radius, and they are either in-phase or out-of-phase with one another. The possible interaction phenomena between two identically sized bubbles are summarized. Depending on their relative distances and phase differences, they can coalesce, jet towards or away from one another, split into smaller bubbles, or 'catapult' away from one another. The 'catapult' effect can be utilized to generated high speed jet in the absence of a solid boundary or shockwave. Also three bubble interactions are highlighted. Complicated phenomena such as bubble forming an elliptical shape and bubble splitting are observed. The BEM simulations provide insight into the physics of the phenomena by providing details such as detailed bubble shape changes (experimental observations are limited by the temporal and spatial resolution), and jet velocity. It is noted that the well-tested BEM code [1,2] utilized here is computationally very efficient as compared to other full-domain methods since only the bubble surface is meshed.

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35

Gibson, T. M., H. Babinsky, and L. C. Squire. "Passive control of shock wave–boundary-layer interactions." Aeronautical Journal 104, no. 1033 (March 2000): 129–40. http://dx.doi.org/10.1017/s000192400002532x.

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Abstract The passive control of a shock wave-boundary-layer interaction involves placing a porous surface beneath the interaction, allowing high pressure air from the flow downstream of the shock wave to recirculate through a plenum chamber into the low pressure flow upstream of the wave. The simple case of a normal shock wave at a Mach number of 1·4 interacting with the turbulent boundary layer on a flat wall is investigated both experimentally and numerically. The experimental investigation made use of holographic interferometry, while the computational section of the investigation made use of a Navier-Stokes code to derive pressure gradients, boundary-layer properties and total pressure losses in the interaction region. It is found that the structure of shock wave-boundary-layer interactions with passive control consists of a leading, oblique shock wave followed by a lambda foot. The oblique wave originates from the upstream end of the porous region, and its strength is determined by the magnitude of the local blowing velocities. The shape of the lambda foot depends on the position of the main shock relative to the control region, resembling an uncontrolled foot when the main shock wave is towards the downstream end of the porosity, but becoming increasingly large as the shock moves upstream and eventually merging with the leading, oblique shock to form a single, large, lambda structure. Improved forms of passive control are suggested based on the findings of this investigation, including the use of passive control systems which incorporate streamwise variations in the level of porosity.

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36

Carroll, Bruce F., and J. Craig Dutton. "Multiple normal shock wave/turbulent boundary-layer interactions." Journal of Propulsion and Power 8, no. 2 (March 1992): 441–48. http://dx.doi.org/10.2514/3.23497.

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37

Hamed, A., S. Shih, and J. Yeuan. "Investigation of shock/turbulent boundary-layer bleed interactions." Journal of Propulsion and Power 10, no. 1 (January 1994): 79–87. http://dx.doi.org/10.2514/3.23714.

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38

Kussoy, M. I., K. C. Horstoman, and C. Horstman. "Hypersonic crossing shock-wave/turbulent-boundary-layer interactions." AIAA Journal 31, no. 12 (December 1993): 2197–203. http://dx.doi.org/10.2514/3.11915.

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39

Macrorie, Michael, and Wayne R. Pauley. "Experimental investigation of convecting vortex/boundary-layer interactions." AIAA Journal 33, no. 8 (August 1995): 1383–90. http://dx.doi.org/10.2514/3.12686.

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40

Brandt, Luca, and H. C. de Lange. "Streak interactions and breakdown in boundary layer flows." Physics of Fluids 20, no. 2 (February 2008): 024107. http://dx.doi.org/10.1063/1.2838594.

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41

Loth, E., and Mark W. Matthys. "Unsteady low Reynolds number shock boundary layer interactions." Physics of Fluids 7, no. 5 (May 1995): 1142–50. http://dx.doi.org/10.1063/1.868555.

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42

Thorpe, S. A. "Interactions between Internal Waves and Boundary Layer Vortices." Journal of Physical Oceanography 27, no. 1 (January 1997): 62–71. http://dx.doi.org/10.1175/1520-0485(1997)027<0062:ibiwab>2.0.co;2.

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43

Hamed, A., and A. Kumar. "Flow Separation in Shock Wave Boundary Layer Interactions." Journal of Engineering for Gas Turbines and Power 116, no. 1 (January 1, 1994): 98–103. http://dx.doi.org/10.1115/1.2906816.

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This work presents an assessment of the experimental data on separated flow in shock wave turbulent boundary layer interactions at hypersonic and supersonic speeds. The data base consists of selected configurations where the only characteristic length in the iteration is the incoming boundary layer thickness. It consists of two-dimensional and axisymmetric interactions in compression corners or cylinder-flares, and externally generated oblique shock interactions with boundary layers over flat plates or cylindrical surfaces. The conditions leading to flow separation and the empirical correlations for incipient separation are reviewed. The effects of Mach number, Reynolds number, surface cooling, and the methods of detecting separation are discussed.

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44

Agostini, Lionel, Michael Leschziner, Jonathan Poggie, Nicholas J. Bisek, and Datta Gaitonde. "Multi-scale interactions in a compressible boundary layer." Journal of Turbulence 18, no. 8 (May 23, 2017): 760–80. http://dx.doi.org/10.1080/14685248.2017.1328108.

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45

Gad-el-Hak, Mohamed. "Boundary Layer Interactions With Compliant Coatings: An Overview." Applied Mechanics Reviews 39, no. 4 (April 1, 1986): 511–24. http://dx.doi.org/10.1115/1.3143723.

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During the past five years, several research programs have been conducted to reexamine the subject of boundary layer interactions with compliant coatings. One of the objectives of the research was to answer the question: Can compliant coatings delay transition and/or significantly reduce turbulence skin friction on bodies at high Reynolds numbers? Several significant developments have been achieved by the many investigators participating in these studies. The purpose of this article is to review the progress in our understanding of compliant coating interactions with laminar, transitional, and turbulent boundary layers. The paper will include some work done prior to the recent five year period and available in the open literature, but will emphasize more recent work, some of which is not as yet published.

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46

John, Bibin, Vinayak N. Kulkarni, and Ganesh Natarajan. "Shock wave boundary layer interactions in hypersonic flows." International Journal of Heat and Mass Transfer 70 (March 2014): 81–90. http://dx.doi.org/10.1016/j.ijheatmasstransfer.2013.10.072.

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47

Dussauge, J. P., and S. Piponniau. "Shock/boundary-layer interactions: Possible sources of unsteadiness." Journal of Fluids and Structures 24, no. 8 (November 2008): 1166–75. http://dx.doi.org/10.1016/j.jfluidstructs.2008.06.003.

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48

Grujicic, Mica, Ramin Yavari, Jennifer Snipes, S. Ramaswami, and Roshdy Barsoum. "All-atom molecular-level computational simulations of planar longitudinal shockwave interactions with polyurea, soda-lime glass and polyurea/glass interfaces." Multidiscipline Modeling in Materials and Structures 10, no. 4 (November 4, 2014): 474–510. http://dx.doi.org/10.1108/mmms-11-2013-0070.

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Purpose – The purpose of this paper is to study the mechanical response of polyurea, soda-lime glass (glass, for short), polyurea/glass/polyurea and glass/polyurea/glass sandwich structures under dynamic-loading conditions involving propagation of planar longitudinal shockwaves. Design/methodology/approach – The problem of shockwave generation, propagation and interaction with material boundaries is investigated using non-equilibrium molecular dynamics. The results obtained are used to construct basic shock Hugoniot relationships associated with the propagation of shockwaves through a homogeneous material (polyurea or glass, in the present case). The fidelity of these relations is established by comparing them with their experimental counterparts, and the observed differences are rationalized in terms of the microstructural changes experienced by the shockwave-swept material. The relationships are subsequently used to predict the outcome of the interactions of shockwaves with polyurea/glass or glass/polyurea material boundaries. Molecular-level simulations are next used to directly analyze the same shockwave/material-boundary interactions. Findings – The molecular-level simulations suggested, and the subsequent detailed microstructural analyses confirmed, the formation of topologically altered interfacial regions, i.e. polyurea/glass and glass/polyurea interphases. Originality/value – To the authors’ knowledge, the present work is a first attempt to analyze, using molecular-level simulation methods, the interaction of shockwaves with material boundaries.

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49

Hilgenfeld, L., P. Cardamone, and L. Fottner. "Boundary layer investigations on a highly loaded transonic compressor cascade with shock/laminar boundary layer interactions." Proceedings of the Institution of Mechanical Engineers, Part A: Journal of Power and Energy 217, no. 4 (January 1, 2003): 349–56. http://dx.doi.org/10.1243/095765003322315405.

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Detailed experimental and numerical investigations of the flowfield and boundary layer on a highly loaded transonic compressor cascade were performed at various Mach and Reynolds numbers representative of real turbomachinery conditions. The emerging shock system interacts with the laminar boundary layer, causing shock-induced separation with turbulent reattachment. Steady two-dimensional calculations have been performed using the Navier—Stokes solver TRACE-U. The flow solver employs a modified version of the one-equation Spalart—Allmaras turbulence model coupled with a transition correlation by Abu-Ghannam/Shaw in the formulation by Drela. The computations reproduce well the experimental results with respect to the profile pressure distribution and the location of the shock system. The transitional behaviour of the boundary layer and the profile losses in the wake are properly predicted as well, except for the highest Mach number tested, where large separated regions appear on the suction side.

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50

Ghosh, Santanu, Jung-Il Choi, and Jack R. Edwards. "Simulation of Shock/Boundary-Layer Interactions with Bleed Using Immersed-Boundary Methods." Journal of Propulsion and Power 26, no. 2 (March 2010): 203–14. http://dx.doi.org/10.2514/1.45297.

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