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Relevant bibliographies by topics / Ramp-induced Shockwave Boundary Layer Interactions

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Academic literature on the topic 'Ramp-induced Shockwave Boundary Layer Interactions'

Author: Grafiati

Published: 6 September 2023

Last updated: 7 September 2023

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Journal articles on the topic "Ramp-induced Shockwave Boundary Layer Interactions"

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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Dissertations / Theses on the topic "Ramp-induced Shockwave Boundary Layer Interactions"

1

Boyer, Nathan Robert. "The Effects of Viscosity and Three-Dimensionality on Shockwave-Induced Panel Flutter." The Ohio State University, 2019. http://rave.ohiolink.edu/etdc/view?acc_num=osu156616766854713.

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Ramji, V. "Experimental Investigations on Ramp-induced Large Separation Bubble in a Hypersonic Flow." Thesis, 2023. https://etd.iisc.ac.in/handle/2005/6136.

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Shockwave Boundary Layer Interactions (SBLIs) are ubiquitous in supersonic/ hypersonic flows and often lead to the separation of the boundary layer. These separations can be broadly classified into small and large separation bubbles based on their lengths compared to the boundary layer thickness. While small bubbles have minimal influence on the outer flow conditions, large bubbles significantly impact the outer flow, causing changes in pressure distribution. These large separation bubbles have been observed to be unsteady and thus can adversely affect the performance of aerodynamic devices. Moreover, these unsteady pressure loads can lead to vibrations and fatigue failure of the structure. Therefore, understanding the flow physics within the separated region and its influence on the outer flow is essential for efficient aerodynamic design. The research gap in the field lies in the accurate prediction of the onset of unsteadiness in Ramp-induced Shockwave Boundary Layer Interactions (R-SBLIs) and the lack of controlled experimental data on unsteady flows with separation occurring at the leading edge. Previous studies have primarily focused on ramp angles below 30°, neglecting higher angles that could lead to detached shock solutions. These gaps in the literature motivate the present study, which aims to investigate the different flow regimes, mechanisms, and sources of low-frequency unsteadiness in R-SBLI. The investigation includes incipient separation, steady separated flow, unsteady separated flow, and the identification of flow topology. A comprehensive approach is proposed for the identification and characterisation of different flow regimes encountered in compression corner flows. The flow regime depends on the pressure ratio imposed by the ramp angle. Shock polar analysis helps identify the nature of shock-shock interactions, which determines the pressure variation along the ramp surface. The oscillation and pulsation modes are identified based on the location of the shear layer impingement on the ramp. A conditional-based algorithm is developed for flow regime identification. This approach provides a systematic understanding of the flow topology for a given set of freestream conditions and test models.

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Greene, Benton Robb. "Control of mean separation in a compression ramp shock boundary layer interaction using pulsed plasma jets." Thesis, 2014. http://hdl.handle.net/2152/25422.

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Pulsed plasma jets (also called "SparkJets'") were investigated for use in controlling the mean separation location induced by shock wave-boundary layer interaction. These synthetic jet actuators are driven by electro-thermal heating from an electrical discharge in a small cavity, which forces the gas in the cavity to exit through a small hole as a high-speed jet. With this method of actuation, pulsed plasma jets can achieve pulsing frequencies on the order of kilohertz, which is on the order of the instability frequency of many lab-scale shock wave-boundary layer interactions (SWBLI). The interaction under investigation was generated by a 20° compression ramp in a Mach 3 flow. The undisturbed boundary layer is transitional with Re[subscript theta] of 5400. Surface oil streak visualization is used in a parametric study to determine the optimum pulsing frequency of the jet, the optimum distance of the jet from the compression corner, and the optimum injection angle of the jets. Three spanwise-oriented arrays of three plasma jets are tested, each with a different pitch and skew angle on the jet exit port. The three injection angles tested were 22° pitch and 45° skew, 20° pitch and 0° skew, and 45° pitch and 0° skew. Jet pulsing frequency is varied between 2 kHz and 4 kHz, corresponding to a Strouhal number based on separation length of 0.012 and 0.023. Particle image velocimetry is used to characterize the effect that the actuators have on the reattached boundary layer profile on the ramp surface. Results show that plasma jets pitched at 20° from the wall, and pulsed at a Strouhal number of 0.018, can reduce the size of an approximate measure of the separation region by up to 40% and increase the integrated momentum in the downstream reattached boundary layer, albeit with a concomitant increase in the shape factor.
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Conference papers on the topic "Ramp-induced Shockwave Boundary Layer Interactions"

1

HORSTMAN, C. "Computation of sharp-fin-induced shockwave/turbulent boundary layer interactions." In 4th Joint Fluid Mechanics, Plasma Dynamics and Lasers Conference. Reston, Virigina: American Institute of Aeronautics and Astronautics, 1986. http://dx.doi.org/10.2514/6.1986-1032.

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Lester, Lauren E., Mark Gragston, Phillip A. Kreth, and John Schmisseur. "Dynamics of Cylinder-Induced Transitional and Turbulent Shockwave-Boundary Layer Interactions on a 6° Cone at Mach 4." In AIAA AVIATION 2021 FORUM. Reston, Virginia: American Institute of Aeronautics and Astronautics, 2021. http://dx.doi.org/10.2514/6.2021-2821.

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Bhamidipati, Keerti K., Daniel A. Reasor, and Crystal L. Pasiliao. "Unstructured Grid Simulations of Transonic Shockwave-Boundary Layer Interaction-Induced Oscillations." In 22nd AIAA Computational Fluid Dynamics Conference. Reston, Virginia: American Institute of Aeronautics and Astronautics, 2015. http://dx.doi.org/10.2514/6.2015-2287.

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Pierce, Adam, Qin Li, Yusi Shih, Frank Lu, and Chaoqun Liu. "Interaction of Microvortex Generator Flow with Ramp-Induced Shock/Boundary-Layer Interactions." In 49th AIAA Aerospace Sciences Meeting including the New Horizons Forum and Aerospace Exposition. Reston, Virigina: American Institute of Aeronautics and Astronautics, 2011. http://dx.doi.org/10.2514/6.2011-32.

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Schöneich, Antonio Giovanni, Thomas J. Whalen, Stuart J. Laurence, Bryson T. Sullivan, Daniel J. Bodony, Maxim Freydin, Earl H. Dowell, Larson J. Stacey, and Gregory M. Buck. "Fluid-Thermal-Structural Interactions in Ramp-Induced Shock-Wave Boundary-Layer Interactions at Mach 6." In AIAA Scitech 2021 Forum. Reston, Virginia: American Institute of Aeronautics and Astronautics, 2021. http://dx.doi.org/10.2514/6.2021-0912.

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Pham, Harry T., Zachary Gianikos, and Venkateswaran Narayanaswamy. "Compression Ramp Induced Shock Wave/Turbulent Boundary Layer Interactions on a Compliant Material." In 2018 AIAA/ASCE/AHS/ASC Structures, Structural Dynamics, and Materials Conference. Reston, Virginia: American Institute of Aeronautics and Astronautics, 2018. http://dx.doi.org/10.2514/6.2018-0095.

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Priebe, Stephan, and M. Pino Martin. "Low-Frequency Unsteadiness in the DNS of a Compression Ramp Shockwave and Turbulent Boundary Layer Interaction." In 48th AIAA Aerospace Sciences Meeting Including the New Horizons Forum and Aerospace Exposition. Reston, Virigina: American Institute of Aeronautics and Astronautics, 2010. http://dx.doi.org/10.2514/6.2010-108.

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Whalen, Thomas J., Richard E. Kennedy, Stuart J. Laurence, Bryson Sullivan, Daniel J. Bodony, and Gregory Buck. "Unsteady Surface and Flowfield Measurements in Ramp-Induced Turbulent and Transitional Shock-Wave Boundary-Layer Interactions at Mach 6." In AIAA Scitech 2019 Forum. Reston, Virginia: American Institute of Aeronautics and Astronautics, 2019. http://dx.doi.org/10.2514/6.2019-1127.

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Smith, Cary D., Phillip A. Kreth, John D. Schmisseur, and Garrett Strickland. "Temperature-Sensitive Paint Measurements of Cylinder-Induced Shockwave-Boundary Layer Interaction on a 6-degree Cone with Laminar Mach 7 Flow." In AIAA SCITECH 2023 Forum. Reston, Virginia: American Institute of Aeronautics and Astronautics, 2023. http://dx.doi.org/10.2514/6.2023-2445.

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