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Academic literature on the topic 'Shock-wave and separation- region interaction'

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Published: 6 September 2023

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Journal articles on the topic "Shock-wave and separation- region interaction"

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Estruch, D., D. G. MacManus, D. P. Richardson, N. J. Lawson, K. P. Garry, and J. L. Stollery. "Experimental study of unsteadiness in supersonic shock-wave/turbulent boundary-layer interactions with separation." Aeronautical Journal 114, no. 1155 (May 2010): 299–308. http://dx.doi.org/10.1017/s0001924000003742.

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AbstractShock-wave/turbulent boundary-layer interactions (SWTBLIs) with separation are known to be inherently unsteady but their physical mechanisms are still not totally understood. An experimental investigation has been performed in a supersonic wind tunnel at a freestream flow Mach number of 2·42. The interaction between a shock wave created by a shock generator (α = 3°, α = 9°, α = 13° and α = 15° deflection angles) and a turbulent boundary layer with thickness δ = 5mm has been studied. High-speed Schlieren visualisations have been obtained and used to measure shock wave unsteadiness by means of digital image processing. In the interactions with separation, the reflected shock’s unsteadiness has been in the order of 102Hz. High-speed wall pressure measurements have also been obtained with fast-response micro-transducers along the interactions. Most of the energy of the incoming turbulent boundary layer is broadband and at high frequencies (>104Hz). An addition of low-frequency (<104Hz) fluctuation energy is found at separation. Along the interaction region, the shock impingement results in an amplification of fluctuation energy due to the increase in pressure. Under the main recirculation region core there is only an increase in high frequency energy (>104Hz). Amplification of lower frequency fluctuation energy (>103Hz) is also observed close to the separation and reattachment regions.
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Mosele, John-Paul, Andreas Gross, and John Slater. "Numerical Investigation of Asymmetric Mach 2.5 Turbulent Shock Wave Boundary Layer Interaction." Aerospace 10, no. 5 (April 29, 2023): 417. http://dx.doi.org/10.3390/aerospace10050417.

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Supersonic shock wave boundary layer interactions are common to inlet flows of supersonic and hypersonic vehicles. This paper reports on wall-resolved implicit large-eddy simulations of a canonical Mach 2.5 turbulent shock wave boundary layer interaction experiment at the NASA Glenn Research Center. The boundary layer upstream of the interaction was nominally axisymmetric and two-dimensional. A conical centerbody with a 16 deg half-angle and a maximum radius of 0.147D of the test section diameter was employed to generate a conical shock wave, where D is the test section diameter. Asymmetric (swept) interactions were obtained by displacing the shock generator away from the test section centerline. The present simulation is for a shock generator displacement of D/6. Results from the asymmetric simulation are compared with results from an earlier simulation of a corresponding axisymmetric interaction. The experimental Reynolds number based on test section diameter was ReD=4×106. For the simulations, the Reynolds number was lowered to ReD=4×105 to keep the computational expense of the simulations within limits. Compared to the axisymmetric interaction, the streamwise extent of the separation varies considerably in the azimuthal direction for the asymmetric interaction. The separation is strongest at the azimuthal location that is closest to the shock generator. The streamwise extent of the separated flow regions is noticeably reduced and substantial crossflow is observed between the locations that are closest and farthest from the shock generator. A Fourier analysis of the unsteady flow data indicates low-frequency content for the separated region that is closest to the shock generator. Away from this region, with increasing sweep angle and cross-flow, the low-frequency content is diminished. A proper orthogonal decomposition captures spanwise coherent structures for the more two-dimensional parts of the interaction.
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Huang, Xin, and David Estruch-Samper. "Low-frequency unsteadiness of swept shock-wave/turbulent-boundary-layer interaction." Journal of Fluid Mechanics 856 (October 11, 2018): 797–821. http://dx.doi.org/10.1017/jfm.2018.735.

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High-speed turbulent boundary-layer separation can lead to severe wall-pressure fluctuations, often extending over a swept shock region. Having noted the shear layer’s influence within axisymmetric step flows, tests go on to experimentally assess the unsteadiness of a canonical swept separation, caused by a slanted $90^{\circ }$-step discontinuity (with varying azimuthal height) over an axisymmetric turbulent boundary layer. Results document an increase in shock pulsation frequency along the swept separation region ($\unicode[STIX]{x1D6EC}\leqslant 30^{\circ }$ sweep angles) – whereby the recirculation enables downstream feedback via the reverse flow – as the local streamwise separation length is reduced. A link between the spanwise variation in the separation shock’s low-frequency instability and the downstream mass ejection rate, as large shear-layer eddies leave the bubble, is sustained. The local entrainment-recharge dynamics of swept separation are thereby duly evaluated.
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Mosele, John-Paul, Andreas Gross, and John Slater. "Numerical Investigation of Mach 2.5 Axisymmetric Turbulent Shock Wave Boundary Layer Interactions." Aerospace 10, no. 2 (February 9, 2023): 159. http://dx.doi.org/10.3390/aerospace10020159.

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Shock wave boundary layer interactions are common to both supersonic and hypersonic inlet flows. Wall-resolved implicit large-eddy simulations of a canonical Mach 2.5 axisymmetric shock wave boundary layer interaction experiment at Glenn Research Center were carried out. A conical shock wave was generated with axisymmetric centerbodies with 16 deg half-angle cone. The centerbody radii were 9.2% and 14.7% of the test section diameter. The conical shock wave interacted with the turbulent boundary layer on the inside of the cylindrical test section. The experimental Reynolds number based on diameter was six million. For the simulations, the Reynolds number was reduced by a factor of 10 to lower the computational expense. The turbulent boundary layer separates for both centerbody radii and the separation is stronger for the larger centerbody radius. Frequency spectra of the spanwise-averaged wall-pressure coefficient reveal low-frequency content at Strouhal numbers based on separation length between 0.02 and 0.05 in the vicinity of the separation shock and mid-frequency content between 0.1 and 0.2 downstream of separation. A proper orthogonal decomposition captures spanwise coherent structures with a Strouhal number of 0.03–0.04 over the interaction region and streamwise coherent structures inside and downstream of the interaction with a Strouhal number of 0.1–0.4.
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Burton, D. M. F., and H. Babinsky. "Corner separation effects for normal shock wave/turbulent boundary layer interactions in rectangular channels." Journal of Fluid Mechanics 707 (August 2, 2012): 287–306. http://dx.doi.org/10.1017/jfm.2012.279.

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AbstractExperiments are conducted to examine the mechanisms behind the coupling between corner separation and separation away from the corner when holding a high-Mach-number ${M}_{\infty } = 1. 5$ normal shock in a rectangular channel. The ensuing shock wave interaction with the boundary layer on the wind tunnel floor and in the corners was studied using laser Doppler anemometry, Pitot probe traverses, pressure sensitive paint and flow visualization. The primary mechanism explaining the link between the corner separation size and the other areas of separation appears to be the generation of compression waves at the corner, which act to smear the adverse pressure gradient imposed upon other parts of the flow. Experimental results indicate that the alteration of the $\lambda $-region, which occurs in the supersonic portion of the shock wave/boundary layer interaction (SBLI), is more important than the generation of any blockage in the subsonic region downstream of the shock wave.
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Chandola, Gaurav, Xin Huang, and David Estruch-Samper. "Highly separated axisymmetric step shock-wave/turbulent-boundary-layer interaction." Journal of Fluid Mechanics 828 (September 6, 2017): 236–70. http://dx.doi.org/10.1017/jfm.2017.522.

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The unsteadiness of a shock-wave/turbulent-boundary-layer interaction induced by an axisymmetric step (cylinder/$90^{\circ }$-disk) is investigated experimentally at Mach 3.9. A large-scale separation of the order of previously reported incoming turbulent superstructures is induced ahead of the step ${\sim}30\unicode[STIX]{x1D6FF}_{o}$ and followed by a downstream separation of ${\sim}10\unicode[STIX]{x1D6FF}_{o}$ behind it, where $\unicode[STIX]{x1D6FF}_{o}$ is the incoming boundary-layer thickness. Narrowband high-frequency instabilities shift gradually to more moderate frequencies along the upstream separation region exhibiting a strong predominance of shear-induced disturbance levels – arising between the outer high-speed flow and the subsonic bubble. Through spectral/time-resolved analysis of this high Reynolds number and large-scale separation, results offer new insights into the shear layer’s inception and evolution (convection, growth and instability) and its influence on interaction unsteadiness.
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Bich Ngoc, Hoang Thi, and Nguyen Manh Hung. "Study of separation phenomenon in transonic flows produced by interaction between shock wave and boundary layer." Vietnam Journal of Mechanics 33, no. 3 (September 8, 2011): 170–81. http://dx.doi.org/10.15625/0866-7136/33/3/210.

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For compressible flows, the transonic state depends on the geometry, Mach number and the incidence. This effect can produce shock wave. Some studies showed that the interaction between shock wave and boundary layer concerns separation phenomenon. Studies in this report demonstrate conditions of separation in transonic flow and that it is not any interaction between shock wave and boundary layer which can cause boundary layer separation. The studies also show that maximum Mach number in the local supersonic region is not a unique factor influencing the separation, and the separation in transonic flows can occur at the incidence of 0\(^{\circ}\). For the calculation of viscous transonic flows, we use Fluent software with serious treatment of application operation based on the physical nature of phenomenon and the technique of numerical treatment. For the calculation of invicid transonic flows, we built a code solving the full potential equation with verification for accuracy. Results calculated from Fluent have been seriously compared with results of present program and published results in order to assure the accuracy of application operation in the domain of investigation. separation in transonic flows; shock wave and boundary layer
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Shahrbabaki, A. Nazarian, M. Bazazzadeh, and R. Khoshkhoo. "Investigation on Supersonic Flow Control Using Nanosecond Dielectric Barrier Discharge Plasma Actuators." International Journal of Aerospace Engineering 2021 (July 14, 2021): 1–14. http://dx.doi.org/10.1155/2021/2047162.

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In this paper, the effects of streamwise Nanosecond Dielectric Barrier Discharge (NS-DBD) actuators on Shock Wave/Boundary Layer Interaction (SWBLI) are investigated in a Mach 2.5 supersonic flow. In this regard, the numerical investigation of NS-DBD plasma actuator effects on unsteady supersonic flow passing a 14° shock wave generator is performed using simulation of Navier-Stokes equations for 3D-flow, unsteady, compressible, and k ‐ ω SST turbulent model. In order to evaluate plasma discharge capabilities, the effects of plasma discharge length on the flow behavior are studied by investigating the flow friction factor, the region of separation bubble formation, velocity, and temperature distribution fields in the SWBLI region. The numerical results showed that plasma discharge increased the temperature of the discharge region and boundary layer temperature in the vicinity of flow separation and consequently reduced the Mach number in the plasma discharge region. Plasma excitation to the separation bubbles shifted the separation region to the upstream around 6 mm, increased SWBLI height, and increased the angle of the separation shock wave. Besides, the investigations on the variations of pressure recovery coefficient illustrated that plasma discharge to the separation bubbles had no impressive effect and decreased pressure recovery coefficient. The numerical results showed that although the NS-DBD plasma actuator was not effective in reducing the separation area in SWBLI, they were capable of shifting the separation shock position upstream. This feature can be used to modify the structure of the shock wave in supersonic intakes in off-design conditions.
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GU, Wenting, Binqian ZHANG, Kun MA, Dong LI, Pengfei LYU, and Jie HAN. "Investigation on the flow mechanism of nacelle airframe interaction for podded blended wing body transport." Xibei Gongye Daxue Xuebao/Journal of Northwestern Polytechnical University 40, no. 2 (April 2022): 352–59. http://dx.doi.org/10.1051/jnwpu/20224020352.

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For the flow interaction between the podded engine and the airframe of blended wing body configuration(BWB), taking the 300 seats class BWB civil transport NPU-BWB-300 designed by Northwestern Polytechnical University as the research object, the influence of the podded engines on the BWB airframe at typical high and low speed conditions were investigated by CFD method, and the airframe-nacelle interference mechanism was revealed. The results indicate that the podded engines mainly affect the high speed performance of BWB, but have little effect on the low speed performance. The flow interaction between the airframe and the nacelle at high speed condition is serious when podded the engines, which leads to strong shock wave and flow separation. The flow mechanism of the above-mentioned interaction is as follows: firstly, the large supersonic region and shock wave on the nacelle external surface interferes with airframe surface flow seriously, which induces shock wave and flow separation; secondly, a convergent-divergent channel is formed between the airframe and the nacelle, resulting in the "throat" effect, which produces shock wave and flow separation.
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LAURENCE, S. J., and R. DEITERDING. "Shock-wave surfing." Journal of Fluid Mechanics 676 (April 6, 2011): 396–431. http://dx.doi.org/10.1017/jfm.2011.57.

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A phenomenon referred to as ‘shock-wave surfing’, in which a body moves in such a way as to follow the shock wave generated by another upstream body, is investigated numerically and analytically. During the surfing process, the downstream body can accumulate a significantly higher lateral velocity than would otherwise be possible. The surfing effect is first investigated for interactions between a sphere and a planar oblique shock. Numerical simulations are performed and a simple analytical model is developed to determine the forces acting on the sphere. A phase-plane description is employed to elucidate features of the system dynamics. The analytical model is then generalised to the more complex situation of aerodynamic interactions between two spheres, and, through comparisons with further computations, is shown to adequately predict the final separation velocity of the surfing sphere in initially touching configurations. Both numerical simulations and a theoretical analysis indicate a strong influence of the sphere radius ratio on the separation process and predict a critical radius ratio that delineates entrainment of the smaller body within the flow region bounded by the larger body's shock from expulsion. Furthermore, it is shown that an earlier scaling law does not accurately describe the separation behaviour. The surfing effect has important implications for meteoroid fragmentation: in particular, a large fraction of the variation in the separation velocity deduced by previous authors from an analysis of terrestrial crater fields can be explained by a combination of surfing and a modest rotation rate of the parent body.
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Dissertations / Theses on the topic "Shock-wave and separation- region interaction"

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Zare, Shahneh Abolghasern. "Investigation of a sub boundary layer vortex generator for the control of separation in boundary layer-shock wave interaction." Thesis, Queen Mary, University of London, 2007. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.485561.

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It is well known that in transonic flows, shQtk waves are fonned over the wings of aircraft. Depending on the strength of the shock wave and the state of the boundary layer, a region of boundary layer separation can occur downstream of the shock. This separation can lead to drastic changes in the flow over the wing which in tum can h~ad to an increase in drag, reduction of lift and buffeting. The aim of this investigation is to assess the potential of sub-boundary layer devices (Sub Boundary Vortex Generators, SBVG) to control a nonnal shock wave/ boundary layer interaction. The experiments have been perfonned at a nominal Mach number of 1.4 and a fre~strcam Reynolds mm1ber of 16x 106 pcr Uii.it length. Detr.ih;d measurements of a fully developed flat plate turbulent boundary layer were used to assess the perfonnance of 8 different SBVG configurations. The SBVG perfonnance was assessed by comparing flow before separation and after the reattachment. Mean flow data such as static and surface total pressure distributions, boundary layer total pressure and velocity profile, surface oil flow visualisation and schlieren method were used in evaluating the perfonnance of SBVGs. The experimental results indicate that the optimum SBVG for the current flow condition is a pair of Tetrahedral Vortex Generator with 30mm length with base and· height dimensions of 3mm by 3mm, which the height is 40% of the boundary layer thickness. The leading and trailing edge of this configuration were 18mm and 3 mm correspondingly.
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Östlund, Jan. "Supersonic flow separation with application to rocket engine nozzles." Doctoral thesis, KTH, Mechanics, 2004. http://urn.kb.se/resolve?urn=urn:nbn:se:kth:diva-3793.

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The increasing demand for higher performance in rocketlaunchers promotes the development of nozzles with higherperformance, which basically is achieved by increasing theexpansion ratio. However, this may lead to flow separation andensuing instationary, asymmetric forces, so-called side-loads,which may present life-limiting constraints on both the nozzleitself and other engine components. Substantial gains can bemade in the engine performance if this problem can be overcome,and hence different methods of separation control have beensuggested. However, none has so far been implemented in fullscale, due to the uncertainties involved in modeling andpredicting the flow phenomena involved.

In the present work the causes of unsteady and unsymmetricalflow separation and resulting side-loads in rocket enginenozzles are investigated. This involves the use of acombination of analytical, numerical and experimental methods,which all are presented in the thesis. A main part of the workis based on sub-scale testing of model nozzles operated withair. Hence, aspects on how to design sub-scale models that areable to capture the relevant physics of full-scale rocketengine nozzles are highlighted. Scaling laws like thosepresented in here are indispensable for extracting side-loadcorrelations from sub-scale tests and applying them tofull-scale nozzles.

Three main types of side-load mechanisms have been observedin the test campaigns, due to: (i) intermittent and randompressure fluctuations, (ii) transition in separation patternand (iii) aeroelastic coupling. All these three types aredescribed and exemplified by test results together withanalysis. A comprehensive, up-to-date review of supersonic flowseparation and side-loads in internal nozzle flows is givenwith an in-depth discussion of different approaches forpredicting the phenomena. This includes methods for predictingshock-induced separation, models for predicting side-loadlevels and aeroelastic coupling effects. Examples are presentedto illustrate the status of various methods, and theiradvantages and shortcomings are discussed.

A major part of the thesis focus on the fundamentalshock-wave turbulent boundary layer interaction (SWTBLI) and aphysical description of the phenomenon is given. Thisdescription is based on theoretical concepts, computationalresults and experimental observation, where, however, emphasisis placed on the rocket-engineering perspective. This workconnects the industrial development of rocket engine nozzles tothe fundamental research of the SWTBLI phenomenon and shows howthese research results can be utilized in real applications.The thesis is concluded with remarks on active and passive flowcontrol in rocket nozzles and directions of futureresearch.

The present work was performed at VAC's Space PropulsionDivision within the framework of European spacecooperation.

Keywords:turbulent, boundary layer, shock wave,interaction, overexpanded,rocket nozzle, flow separation,control, side-load, experiments, models, review.

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Östlund, Jan. "Flow Processes in Rocket Engine Nozzles with Focus on Flow Separation and Side-Loads." Licentiate thesis, KTH, Mechanics, 2002. http://urn.kb.se/resolve?urn=urn:nbn:se:kth:diva-1452.

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Kowalczyk, Piotr Jozef. "Validation and application of advanced soil constitutive models in numerical modelling of soil and soil-structure interaction under seismic loading." Doctoral thesis, Università degli studi di Trento, 2020. http://hdl.handle.net/11572/275675.

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This thesis presents validation and application of advanced soil constitutive models in cases of seismic loading conditions. Firstly, results of three advanced soil constitutive models are compared with examples of shear stack experimental data for free field response in dry sand for shear and compression wave propagation. Higher harmonic generation in acceleration records, observed in experimental works, is shown to be possibly the result of soil nonlinearity and fast elastic unloading waves. This finding is shown to have high importance on structural response, real earthquake records and reliability of conventionally employed numerical tools. Finally, short study of free field response in saturated soil reveals similar findings on higher harmonic generation. Secondly, two advanced soil constitutive models are used, and their performance is assessed based on examples of experimental data on piles in dry sand in order to validate the ability of the constitutive models to simulate seismic soil-structure interaction. The validation includes various experimental configurations and input motions. The discussion on the results focuses on constitutive and numerical modelling aspects. Some improvements in the formulations of the models are suggested based on the detailed investigation. Finally, the application of one of the advanced soil constitutive models is shown in regard to temporary natural frequency wandering observed in structures subjected to earthquakes. Results show that pore pressure generated during seismic events causes changes in soil stiffness, thus affecting the natural frequency of the structure during and just after the seismic event. Parametric studies present how soil permeability, soil density, input motion or a type of structure may affect the structural natural frequency and time for its return to the initial value. In addition, a time history with an aftershock is analysed to investigate the difference in structural response during the earthquake and the aftershock.
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Kowalczyk, Piotr Jozef. "Validation and application of advanced soil constitutive models in numerical modelling of soil and soil-structure interaction under seismic loading." Doctoral thesis, Università degli studi di Trento, 2020. http://hdl.handle.net/11572/275675.

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This thesis presents validation and application of advanced soil constitutive models in cases of seismic loading conditions. Firstly, results of three advanced soil constitutive models are compared with examples of shear stack experimental data for free field response in dry sand for shear and compression wave propagation. Higher harmonic generation in acceleration records, observed in experimental works, is shown to be possibly the result of soil nonlinearity and fast elastic unloading waves. This finding is shown to have high importance on structural response, real earthquake records and reliability of conventionally employed numerical tools. Finally, short study of free field response in saturated soil reveals similar findings on higher harmonic generation. Secondly, two advanced soil constitutive models are used, and their performance is assessed based on examples of experimental data on piles in dry sand in order to validate the ability of the constitutive models to simulate seismic soil-structure interaction. The validation includes various experimental configurations and input motions. The discussion on the results focuses on constitutive and numerical modelling aspects. Some improvements in the formulations of the models are suggested based on the detailed investigation. Finally, the application of one of the advanced soil constitutive models is shown in regard to temporary natural frequency wandering observed in structures subjected to earthquakes. Results show that pore pressure generated during seismic events causes changes in soil stiffness, thus affecting the natural frequency of the structure during and just after the seismic event. Parametric studies present how soil permeability, soil density, input motion or a type of structure may affect the structural natural frequency and time for its return to the initial value. In addition, a time history with an aftershock is analysed to investigate the difference in structural response during the earthquake and the aftershock.
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Diop, Moussa. "Transition à la turbulence en écoulements compressibles décollés." Thesis, Aix-Marseille, 2017. http://www.theses.fr/2017AIXM0473/document.

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Les recherches sur les instationnarités des Interactions Ondes de Choc Couches Limites (IOCCL) turbulentes ont permis une description détaillée de celles-ci tant expérimentalement que numériquement . Ceci a conduit à plusieurs schémas susceptibles d'expliquer les respirations à basses fréquences observées dans de tels écoulements. Les configurations avec des conditions amont laminaires ou transitionnelles ont été moins étudiées.Dans le cadre du programme Européen TFAST, un important effort a été mené afin de développer des dispositifs expérimentaux, conjointement à des simulations numériques, permettant une étude détaillée de ces configurations. Dans le cadre de cette thèse, on a mis en place une configuration de réflexion d'onde de choc sur une couche limite laminaire pour un nombre de Mach de 1.68. L'utilisation des métrologies classiques (Anémométrie Laser Doppler, Anémométrie Fil Chaud), adaptées à ces conditions expérimentales particulières, a permis de décrire les propriétés spatio-temporelles de ces écoulements. Le champ moyen a été caractérisé et comparé aux théories classique et aux résultats obtenus dans différentes souffleries.Un schéma décrivant le mécanisme de transition à la turbulence au sein de l'interaction a été développé. Sa sensibilité aux conditions amont a été étudiée en plaçant des perturbations en amont de l'interaction. Dans tous les cas, des instationnarités convectives (haute fréquence) et stationnaires (basse fréquence) ont été observées et comparées à celles existantes pour les configurations amont turbulentes. Une gamme intermédiaire d'instationnarités convectives (moyenne fréquence) a été mise en évidence et caractérisée
Research dedicated to the study of the unsteadiness of turbulent Shock Wave Boundary Layer Interaction (SWBLI) has allowed a detailed description of this kind of interaction both experimentally and numerically. Several scenario were proposed to explain the low frequency unsteadiness observed in separated SWBLI. Nevertheless, the literature on this kind of flow involving either upstream laminar or transitional conditions is quite reduce. Within the framework of the European TFAST program, an important effort was made to develop experimental devices, in conjunction with numerical simulations, allowing a detailed study of these laminar or transitional configurations. In particular, within the framework of this thesis, a shock wave reflection configuration on a laminar boundary layer was set-up, with a nominal free stream Mach number of 1.68. Using classical metrology (Laser Doppler Anemometry, Hot WireAnemometry) that have been adapted to these particular experimental conditions, we have been able to describe the spatio-temporal properties of the interaction. The mean field has been characterized and compared with the classical theories and the results obtained in other configurations.A model describing the transition mechanisms to turbulence within the interaction has been developed. Its sensitivity to upstream conditions was studied by placing perturbations upstream of the interaction. In all cases, convective (high frequency) and stationary (low frequency) unsteadiness were observed and compared with those existing for upstream turbulent configurations. An intermediate range of convective unsteadiness (medium frequency) has been demonstrated and characterized
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Kumara, Akshaya G. "Small-amplitude Oscillations in Hypersonic Double-cone Flow." Thesis, 2023. https://etd.iisc.ac.in/handle/2005/6030.

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Unsteady compressible flows typically pose problems with rich dynamics. The broad concern of this thesis is the shock wave unsteadiness that arises in external high-speed flow over a double-cone body. This unsteadiness is driven by complex interaction between the shock wave and the region of high shear and separation in the external flow. It is well known that the canonical double-cone problem exhibits two different classes of unsteadiness, labeled “pulsations” and “oscillations.” The former is characterized by unsteady shock wave motion over large spatial scales, whereas in the latter the nature of unsteadiness is distinct and occurs at a relatively smaller scale. The detailed mechanisms that sustain pulsations and oscillations are yet to be completely understood. In the present work, experiments were performed in the Roddam Narasimha Hypersonic Wind Tunnel (RNHWT) at Mach 6 to carefully investigate the phenomena of oscillations. Time-resolved schlieren and wall pressure data were obtained for various double-cone models with the second cone angle fixed at 90◦, while the first cone angle and ratio of the slant lengths were varied as parameters. Schlieren data revealed two dissimilar types, or modes, of flow oscillations. Spectral proper orthogonal decomposition (SPOD) analysis performed on experimental data showed the existence of a dominant time scale for the oscillations, and also provided the associated low-rank dynamics. The two different oscillation modes are found to exhibit distinct Strouhal number scaling. Given the direct dependence of shock strength on the flow Mach number (M ), the boundaries of unsteady flow states are expected to show slight changes with M. However, qualitative flow features and the underlying mechanisms that drive unsteadiness are expected to remain the same. Overall, this work reveals new flow behavior and furthers our understanding of the double-cone flow.
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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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Narayanaswamy, Venkateswa. "Investigation of a pulsed-plasma jet for separation shock/boundary layer interaction control." Thesis, 2010. http://hdl.handle.net/2152/ETD-UT-2010-05-1400.

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A pulsed-plasma jet (called a "spark-jet" by other researchers), is a high-speed synthetic jet that is generated by striking an electrical discharge in a small cavity. The gas in the cavity pressurizes owing to the heating and is allowed to escape through a small orifice. A series of experiments were conducted to determine the characteristics of the pulsed-plasma jet issuing into stagnant air at a pressure of 45 Torr. These results show that typical jet exit velocities of about 250 m/s can be induced with discharge energies of about 30 mJ per jet. Furthermore, the maximum pulsing frequency was found to be about 5 kHz, because above this frequency the jet begins to misfire. The misfiring appears to be due to the finite time it takes for the cavity to be recharged with ambient air between discharge pulses. The velocity at the exit of the jet is found to be primarily dependent on the discharge current and independent of other discharge parameters such as cavity volume and orifice diameter. Temperature measurements are made using optical emission spectroscopy and reveal the presence of considerable non-equilibrium between rotational and vibrational modes. The gas heating efficiency was found to be 10% and this parameter is shown to have a direct effect on the plasma jet velocity. These results indicate that the pulsed-plasma jet creates a sufficiently strong flow perturbation that is holds great promise as a supersonic flow actuator. An experimental study is conducted to characterize the performance of a pulsed-plasma jet for potential use in supersonic flow control applications. To obtain an estimate of the relative strength of the pulsed-plasma jet, the jet is injected normally into a Mach 3 cross-flow and the penetration distance is measured by using schlieren imaging. These measurements show that the jet penetrates 1.5 [delta], where [delta] is the boundary layer thickness, into the cross-flow and the jet-to-crossflow momentum flux ratio is estimated to be 0.6. An array of pulsed-plasma jets was issued from different locations upstream of a 30-degree compression ramp in a Mach 3 flow. Furthermore, two different jet configurations were used: normal injection and pitched and skewed injection. The pitched and skewed configuration was used to see if the jets could act as high-bandwidth pulsed vortex generators. The interaction between the jets and the separation shock was studied using phase-locked schlieren imaging. Results show that the plasma jets cause a significant disturbance to the separation shock and clearly influence its unsteadiness. While all plasma jet configurations tested caused an upstream motion of the separation shock, pitched and skewed plasma jets caused an initial downstream shock motion before the upstream motion, demonstrating the potential use of these plasma jets as vortex generator jets. The effect of the plasma jet array on the separation shock unsteadiness is studied in a time-resolved manner by using 10 kHz schlieren imaging and fast-response wall pressure measurements. An array of three pulsed-plasma jets, in a pitched and skewed configuration, is used to force the unsteady motion of the interaction formed by a 24° compression ramp in a Mach 3 flow. The Reynolds number of the incoming boundary layer is Re[theta]=3300. Results show that when the pulsed jet array is placed upstream of the interaction, the jets cause the separation shock to move in a quasi-periodic manner, i.e., nearly in sync with the pulsing cycle. As the jet fluid convects across the separation shock, the shock responds by moving upstream, which is primarily due to the presence of hot gas and hence the lower effective Mach number of the incoming flow. Once the hot gases pass through the interaction, the separation shock recovers by moving downstream, and this recovery velocity is approximately 1% to 3% of the free stream velocity. With forcing, the low-frequency energy content of the pressure fluctuations at a given location under the intermittent region decreases significantly. This is believed to be a result of an increase in the mean scale of the interaction under forced conditions. Pulsed-jet injection are also employed within the separation bubble, but negligible changes to the separation shock motion were observed. These results indicate that influencing the dynamics of this compression ramp interaction is much more effective by placing the actuator in the upstream boundary layer.
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Books on the topic "Shock-wave and separation- region interaction"

1

Hamed, A. Flow separation in shock wave boundary layer interactions at hypersonic speeds. Washington, D.C: National Aeronautics and Space Administration, Office of Management, Scientific and Technical Information Division, 1990.

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W, Barter J., and United States. National Aeronautics and Space Administration., eds. Control & reduction of unsteady pressure loads in separated shock wave turbulent boundary layer interaction: Final report on NASA grant NAG 1-1471 for the period 01/09/93 through 01/01/95. Austin, Tex: Center for Aerodynamics Research, Dept. of Aerospace Engineering & Engineering Mechanics, University of Texas at Austin, 1995.

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W, Barter J., and United States. National Aeronautics and Space Administration., eds. Control & reduction of unsteady pressure loads in separated shock wave turbulent boundary layer interaction: Final report on NASA grant NAG 1-1471 for the period 01/09/93 through 01/01/95. Austin, Tex: Center for Aerodynamics Research, Dept. of Aerospace Engineering & Engineering Mechanics, University of Texas at Austin, 1995.

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W, Barter J., and United States. National Aeronautics and Space Administration., eds. Control & reduction of unsteady pressure loads in separated shock wave turbulent boundary layer interaction: Final report on NASA grant NAG 1-1471 for the period 01/09/93 through 01/01/95. Austin, Tex: Center for Aerodynamics Research, Dept. of Aerospace Engineering & Engineering Mechanics, University of Texas at Austin, 1995.

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D, Saunders J., and United States. National Aeronautics and Space Administration., eds. 3D Navier-Stokes analysis of a Mach 2.68 bifurcated rectangular mixed-compression inlet. [Washington, DC]: National Aeronautics and Space Administration, 1995.

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D, Saunders J., and United States. National Aeronautics and Space Administration., eds. 3D Navier-Stokes analysis of a Mach 2.68 bifurcated rectangular mixed-compression inlet. [Washington, DC]: National Aeronautics and Space Administration, 1995.

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D, Saunders J., and United States. National Aeronautics and Space Administration., eds. 3D Navier-Stokes analysis of a Mach 2.68 bifurcated rectangular mixed-compression inlet. [Washington, DC]: National Aeronautics and Space Administration, 1995.

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D, Saunders J., and United States. National Aeronautics and Space Administration., eds. 3D Navier-Stokes analysis of a Mach 2.68 bifurcated rectangular mixed-compression inlet. [Washington, DC]: National Aeronautics and Space Administration, 1995.

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Simulation of glancing shock wave and boundary layer interaction. Moffett Field, Calif: National Aeronautics and Space Administration, Ames Research Center, 1989.

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Zeitlin, Vladimir. Geostrophic Adjustment and Wave–Vortex (Non)Interaction. Oxford University Press, 2018. http://dx.doi.org/10.1093/oso/9780198804338.003.0008.

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The fundamental process of geostrophic adjustment is treated by the method of multi-scale asymptotic expansions in Rossby number and fast-time averaging (which is explained), first in the barotropic one-layer case, and then in the baroclinic two-layer case. Together with the standard quasi-geostrophic regime of parameters, the frontal (or semi-) geostrophic regime is considered. Dynamical separation of slow and fast motions is demonstrated in both regimes. The former obey quasi-geostrophic or frontal-geostrophic equations, thus providing formal justification of the heuristic derivation of Chapter 5. Fast motions are inertia-gravity waves in quasi-geostrophic case, and inertial oscillations in the frontal-geostrophic case. Geostrophic adjustment is also considered in the presence of coastal, topographic, and equatorial wave-guides, and, again, separation of fast and slow motions is demonstrated, the latter now including long Kelvin waves in the first case, long topographic waves in the second case, and long Kelvin and Rossby waves in the third case.
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Book chapters on the topic "Shock-wave and separation- region interaction"

1

Georgievskiy, P. Y., and V. A. Levin. "Front separation regions for blunt and streamlined bodies initiated by temperature wake – bow shock wave interaction." In Shock Waves, 1273–78. Berlin, Heidelberg: Springer Berlin Heidelberg, 2009. http://dx.doi.org/10.1007/978-3-540-85181-3_77.

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Aso, Shigeru, Keiichi Karashima, Kiyoshi Sato, Satoshi Okuyama, and Shozo Maekawa. "Flow Visualization of Secondary Separation and Oscillating Shock Waves in Three-Dimensional Shock Waves-Turbulent Boundary Layer Interaction Region." In Flow Visualization VI, 607–11. Berlin, Heidelberg: Springer Berlin Heidelberg, 1992. http://dx.doi.org/10.1007/978-3-642-84824-7_107.

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Debiève, J. F., and P. Dupont. "Dependence Between Shock and Separation Bubble in a Shock Wave Boundary Layer Interaction." In IUTAM Symposium on Unsteady Separated Flows and their Control, 331–41. Dordrecht: Springer Netherlands, 2009. http://dx.doi.org/10.1007/978-1-4020-9898-7_28.

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Maekawa, Syozo, Shigeru Aso, Shigehide Nakao, Kazuo Arashi, Kenji Tomioka, and Hiroyuki Yamao. "Aerodynamic Heating in Three-Dimensional Bow Shock Wave/Turbulent Boundary Layer Interaction Region." In Shock Waves @ Marseille I, 133–38. Berlin, Heidelberg: Springer Berlin Heidelberg, 1995. http://dx.doi.org/10.1007/978-3-642-78829-1_21.

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Huang, X., G. Chandola, and D. Estruch-Samper. "Unsteady Separation Shock Dynamics in a Mach 4 Shock-Wave/Turbulent Boundary Layer Interaction." In 31st International Symposium on Shock Waves 1, 1007–13. Cham: Springer International Publishing, 2019. http://dx.doi.org/10.1007/978-3-319-91020-8_121.

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Dallmann, U., and P. Doerffer. "Three-Dimensional Flow Separation caused by Normal Shock-Wave / Turbulent Boundary-Layer Interaction." In Symposium Transsonicum III, 429–38. Berlin, Heidelberg: Springer Berlin Heidelberg, 1989. http://dx.doi.org/10.1007/978-3-642-83584-1_35.

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Chandola, G., X. Huang, and D. Estruch-Samper. "Experimental Study on the Unsteadiness of an Axisymmetric Shock-Wave/Turbulent-Boundary-Layer Interaction with Separation." In 31st International Symposium on Shock Waves 1, 1067–73. Cham: Springer International Publishing, 2019. http://dx.doi.org/10.1007/978-3-319-91020-8_128.

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Caballero, N. "Drag Reduction in Airfoils Using Control Devices in the Shock Wave-Boundary Layer Interaction Region." In Aerodynamic Drag Reduction Technologies, 377–84. Berlin, Heidelberg: Springer Berlin Heidelberg, 2001. http://dx.doi.org/10.1007/978-3-540-45359-8_40.

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Daub, Dennis, Sebastian Willems, Burkard Esser, and Ali Gülhan. "Experiments on Aerothermal Supersonic Fluid-Structure Interaction." In Notes on Numerical Fluid Mechanics and Multidisciplinary Design, 323–39. Cham: Springer International Publishing, 2020. http://dx.doi.org/10.1007/978-3-030-53847-7_21.

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Abstract Mastering aerothermal fluid-structure interaction (FSI) is crucial for the efficient and reliable design of future (reusable) launch vehicles. However, capabilities in this area are still quite limited. To address this issue, a multidisciplinary experimental and numerical study of such problems was conducted within SFB TRR 40. Our work during the last funding period was focused on studying the effects of moderate and high thermal loads. This paper provides an overview of our experiments on FSI including structural dynamics and thermal effects for configurations in two different flow regimes. The first setup was designed to study the combined effects of thermal and pressure loads. We investigated a range of conditions including shock-wave/boundary-layer interaction (SWBLI) with various incident shock angles leading to, in some cases, large flow separation with high amplitude temperature dependent panel oscillations. The respective aerothermal loads were studied in detail using a rigid reference panel. The second setup allowed us to study the effects of severe heating leading to plastic deformation of the structure. We obtained severe localized heating resulting in partly plastic deformations of more than 12 times the panel thickness. Furthermore, the effects of repeated load cycles were studied.
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Renane, Rachid, Rachid Allouche, Oumaima Zmit, and Bouchra Bouchama. "Aero Heating Optimization of a Hypersonic Thermochemical Non-Equilibrium Flow around Blunt Body by Application of Opposing Jet and Blunt Spike." In Hypersonic Vehicles - Applications, Recent Advances, and Perspectives [Working Title]. IntechOpen, 2022. http://dx.doi.org/10.5772/intechopen.101659.

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The goal of this work is to give optimum aerothermal solutions for thermal protection of the nose wall of space shuttles during atmospheric reentry, where the air flow is hypersonic, nonequilibrium reactive flow (vibrational and chemical) behind detached shock waves, it’s governed by Navier–Stokes equations with chemical reaction source terms, and modelled using five species (N2, O2, NO, N, O) and Zeldovich chemical scheme with five reactions. This study which simulates the flow using the software Fluent v.19 focuses on the comparison between three protection techniques based on the repulsion of the shock wave, the first is geometric, it consists in introducing a spike that makes the right shock move away from the nose of the shuttle, this allows the endothermic physicochemical processes of dissociation and ionization to absorb heat, the second technique is based on an opposite jet configuration in the frontal region of the nose, this jet allows to push the strong shock, and consequently reduce the heat released, the last technique is the assembly of the two previous techniques; Jets nearby the spike noses were set up in front of the blunt body to reconfigure the flow field and reduce aerodynamic overheating. The opposing jet model reduces the heat at the nose by 12.08% compared to the spike model and by 20.36% compared to the spike jet model. The flow field reconfiguration was the most important factor in heat reduction, according to the quantitative analysis, a combination parameter was given as the main criterion for designing spiked bodies with opposing jets for the goal of heat reduction based on the locations of the reattached shock and its interaction with the conical shock. The results obtained are in good agreement with the specialized literature.
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Conference papers on the topic "Shock-wave and separation- region interaction"

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Jian, Liu, Duan Wenhua, Zhang Liangji, and Qiao Weiyang. "Effect of Suction Side Jet on the Shock Wave Boundary Layer Interaction in Transonic Turbine." In ASME Turbo Expo 2020: Turbomachinery Technical Conference and Exposition. American Society of Mechanical Engineers, 2020. http://dx.doi.org/10.1115/gt2020-16256.

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Abstract In this paper, the effect of round jet with inclination angle 135° upstream the throat on the suction surface on shock wave boundary-layer interaction was investigated in a transonic turbine cascade, and the vortical structures near the jet region were analyzed. Owing to locally high concave curvature on the pressure side profile, the double shock wave structure was obtained in the turbine passage near the pressure side trailing edge. The first incident shock does not induce the boundary layer separation. The second strong incident shock transmits from the trailing edge of the pressure side and reaches the suction side of the adjacent blade. Strong interaction between the suction side boundary layer and incident shock wave exists in this region, and the separation bubble appears in the no jet case. The complex shock wave system and corresponding flow characters are analyzed. Due to the complex vortical structures on the blade suction surface with suction side jet, the pressure distribution on the suction side changes, and the shock wave system in the transonic turbine passage is rearranged, thereby influencing the shock wave boundary layer interaction. The separation onset decays with the suction side jet, and it keeps move downstream with increasing jet velocity. Length of the separation bubble is significantly reduced with suction side jet. However, when the jet velocity is beyond a certain value, the effect of suction side jet will not improve. The complex vortical structures with suction side jet will reenergizing to the low momentum fluid within boundary layer, and the mean velocity profiles in the boundary layer near the shock wave boundary layer interaction religion with suction side jet are more solid than the no jet case, which infers stronger resistance to flow separation. Complicated vortical structures exist near jet region, the Kelvin–Helmholtz instabilities of the shear layer of the jet flow and its coherent structures dominate the unsteadiness of the suction surface. The incident shock wave enhances the pressure fluctuation in the SBLI region, whereas the effect concentrates only on the first harmonic of the K-H instability but not higher frequencies.
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Shi, Ke, and Song Fu. "Study of Shock/Blade Tip Leakage Vortex/Boundary Layer Interaction in a Transonic Rotor With IDDES Method." In ASME Turbo Expo 2013: Turbine Technical Conference and Exposition. American Society of Mechanical Engineers, 2013. http://dx.doi.org/10.1115/gt2013-95252.

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In the present study, Improved Delayed Detached Eddy Simulation (IDDES) based on k-ω-SST turbulence model is applied to study the unsteady phenomenon in a transonic compressor rotor. Particular emphasis is on the understanding of the complex underlying mechanisms for the flow unsteadiness caused by the interaction of passage shock, blade tip leakage vortex (BTLV) and the blade boundary layer. The sources of the significant unsteadiness of the flow are shown. At the lower span height, where the BTLV is far away, the shock wave ahead of the blade leading edge impinges on the suction surface boundary layer of the adjacent blade, causing the shock wave/boundary layer interaction (SWBLI). Boundary layer thickness grows, while flow separates after the interaction. Predicted by IDDES calculation, this shock-induced separation exists as a separation bubble. The flow reattaches very soon after separation. At the near tip region, the shock wave surface deforms due to the strong interaction between the shock and the BTLV. Oscillation of the shock wave surface near the vortex core infers an unsteady contend between the shock and the vortex. Iso-surfaces of the Q parameter are applied to identify the vortex and its structure. Normally, the vortex breakdown in the rotor passage will lead to stall. However, in the present transonic case, the vortex breakdown was observed even at the near peak efficiency point. While the mass flow rate decreases, the shock waves formed ahead of the rotor blade leading edge were pushed upstream, causing earlier casing wall boundary layer separation. Upstream moving behavior of the shock is considered a new stall process.
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Schreiber, H. A., and H. Starken. "An Investigation of a Strong Shock-Wave Turbulent Boundary Layer Interaction in a Supersonic Compressor Cascade." In ASME 1991 International Gas Turbine and Aeroengine Congress and Exposition. American Society of Mechanical Engineers, 1991. http://dx.doi.org/10.1115/91-gt-092.

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Experiments have been performed in a Supersonic cascade facility to elucidate the fluid dynamic phenomena and loss mechanism of a strong shock-wave turbulent boundary layer interaction in a compressor cascade. The cascade geometry is typical for a transonic fan tip section that operates with a relative inlet Mach number of 1.5, a flow turning of about 3 degrees, and a static pressure ratio of 2.15. The strong oblique and partly normal blade passage shock-wave with a pre-shock Mach number level of 1.42 to 1.52 induces a turbulent boundary layer separation on the blade suction surface. Freestream Reynolds number based on chord length was about 2.7×106. Cascade overall performance, blade surface pressure distributions, Schlieren photographs, and surface visualisations are presented. Detailed Mach number and flow direction profiles of the interaction region (lambda shock) and the corresponding boundary layer have been determined using a Laser-2-Focus anemometer. The obtained results indicated that the axial blade passage stream sheet contraction (axial velocity density ratio) has a significant influence on the mechanism of strong interaction and the resulting total pressure losses.
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Saito, Seishiro, Masato Furukawa, Kazutoyo Yamada, Keisuke Watanabe, Akinori Matsuoka, and Naoyuki Niwa. "Mechanisms and Quantitative Evaluation of Flow Loss Generation in a Multi-Stage Transonic Axial Compressor." In ASME Turbo Expo 2019: Turbomachinery Technical Conference and Exposition. American Society of Mechanical Engineers, 2019. http://dx.doi.org/10.1115/gt2019-90439.

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Abstract Flow structure and flow loss generation in a transonic axial compressor has been numerically investigated by using a large-scale detached eddy simulation (DES). The data mining techniques, which include a vortex identification based on the critical point theory and a limiting streamline visualization with the line integral convolution (LIC) method, were applied to the DES result in order to analyze the complicated flow field in compressor. The flow loss in unsteady flow field was evaluated by entropy production rate, and the loss mechanism and the loss amount of each flow phenomenon were investigated for the first rotor and the first stator. In the first rotor, a shock-induced separation is caused by the detached shock wave and the passage shock wave. On the hub side, a hub-corner separation occurs due to the secondary flow on the hub surface, and a hub-corner separation vortex is clearly formed. The flow loss is mainly caused by the blade boundary layer and wake, and the loss due to the shock wave is very small, only about 1 percent of the total loss amount in the first rotor. However, the shock/boundary layer interaction causes an additional loss in the blade boundary layer and the wake, which amount reaches to about 30 percent of the total. In the first stator, the hub-corner separation occurs on the suction side. Although only one hub-corner separation vortex is formed in the averaged flow field, the hub-corner separation vortex is generated in multiple pieces and those pieces interfere with each other in an instantaneous flow field. The hub-corner separation generates huge loss over a wide range, however, the loss generation around the hub-corner separation vortex is not so large, and the flow loss is mainly produced in the shear layer between the mainstream region and the separation region. The main factors of loss generation are the boundary layer, wake and hub-corner separation, which account for about 80 percent of the total loss amount in the first stator.
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Bell, Ralf M., and Leonhard Fottner. "Investigations of Shock/Boundary-Layer Interaction in a Highly Loaded Compressor Cascade." In ASME 1995 International Gas Turbine and Aeroengine Congress and Exposition. American Society of Mechanical Engineers, 1995. http://dx.doi.org/10.1115/95-gt-084.

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Experimental investigations of the shock/boundary-layer interaction were carried out in a highly loaded compressor cascade under realistic turbomachinery conditions in order to improve the accuracy of semi-empirical flow and loss prediction methods. Different shock positions and strengths were obtained by variations of inlet flow angle and inlet Mach number. The free stream turbulence intensity, depending on the inlet Mach number, changed between 4% and 8%. The influence of the inlet Reynolds number based on blade chord is also examined for two different values (Re1=450000, 900000). Schlieren pictures of the transonic cascade flow reveal an unsteady flow behavior with different shock configurations, depending on the pre-shock Mach number. Wake distributions and boundary-layer measurements with the Laser two-focus velocimetry show that the increase of total pressure loss with increasing inlet Mach number is mainly due to the shock/boundary-layer interaction. The shock interaction with a laminar/transitional boundary-layer causes a wide streamwise pressure diffusion, clearly shown by profile pressure distributions. This has a strong influence on the flow outside of the boundary-layer presented by a quantitative Schlieren image. The transition process, investigated with the analysis of thin-film signals, is induced by the shock-wave and occurs above a separated-flow region. At the higher Reynolds number a shock-induced transition takes place without separation.
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Czerwinska, J., P. Doerffer, and F. Magagnato. "Bifurcation of Shock Induced Separation Structures." In ASME 2001 International Mechanical Engineering Congress and Exposition. American Society of Mechanical Engineers, 2001. http://dx.doi.org/10.1115/imece2001/fed-24928.

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Abstract The scope of this paper is to present bifurcation phenomenon which occur in the transonic flow of a curved channel. The channel geometry is nominally two-dimensional. Due to changes in the third dimension, which seems to be a bifurcation parameter, shock induced separation structures significantly change their behavior. The objective of this paper is to show the phenomenon as well as to present models and the explanation of the physical process causing bifurcation in shock induced separation structures. The interest for consideration of the presented flow has its origin in turbomachinery. The channel flow is a simplification, which allows the study of basic phenomenon of boundary layer shock wave interactions. Shock induced separation structures occur on the walls of the channel. For a specific geometry taken into account here, such structures appear only on the bottom and side walls. The changes in their character occur due to increase in the distance between side walls. Two channel width cases (50 mm and 150mm) have been particularly well studied. The cases will be later referred to as either narrow or wide passages respectively. For the first one the separation structures were wide and reach far down stream, while for a second case separation was much smaller. For both cases experiments and 3-D Navier-Stokes calculations with turbulence modeling were performed. The separation structures can be seen also as sets of focal and saddle points, which change their type (Takens-Bogdanov bifurcation) with the changes of the distance between side walls. The description of the flow structures and their bifurcation has established a base for one aspect. Another consideration is global bifurcation. The question is how to predict what structures should be expected for any given width between the two cases. Further, how does one predict if small changes of the side wall distance cause small changes in separation structures or if they have a more critical behavior. The experiment and also 3-D calculations are very expensive and time consuming and they are not able to answer this in an easy way. Hence, a theoretical model is presented, which can describe global bifurcation of flow structures. In experiments as well as in the full 3-D Navier-Stokes calculations the third dimension of the separation region was significantly smaller and depends on the Reynolds number, as does the height of the λ-foot. This has given the idea to model only a separation region in a way similar to the Hele-Shaw flow. As a result a simpler two-dimensional equation was obtained. The solution can be achieved in an easier and more efficient way than the calculation of unsimplified cases. This gives possibility of an estimation of the flow for the other values of the channel width. The proposed paper presents a phenomenon observed in shock induced separation structures in transonic turbulent flow by means of experiments and numerical simulations. Further, it attempts to explain the physical processes involved in this phenomenon by deriving a theoretical model.
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Priebe, Stephan, Daniel Wilkin, Andy Breeze-Stringfellow, Arash Mousavi, Rathakrishnan Bhaskaran, and Luke d’Aquila. "Large Eddy Simulations of a Transonic Airfoil Cascade." In ASME Turbo Expo 2022: Turbomachinery Technical Conference and Exposition. American Society of Mechanical Engineers, 2022. http://dx.doi.org/10.1115/gt2022-80683.

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Abstract The tip region of transonic blades in turbomachinery involves complex flow physics including shock wave/boundary layer interactions (SBLI). SBLI can lead to flow separation, transition from laminar to turbulent flow, and unsteadiness, which can affect the overall performance of the blade. In this paper, we present wall-resolved large eddy simulations (LES) of a transonic rotating cascade that is modeled after the tip region of a transonic diffusing blade. The calculations were performed using GENESIS, a high-order unstructured large eddy simulation solver. The convergence of the LES solution is assessed by varying the polynomial order of the solution from low to high order. LES simulations for a total of five operating conditions are presented, which cover the range of operation from unique incidence low operating line to stall and the associated shock wave/boundary layer interaction physics. The overall aerodynamics of the transonic passage airfoil are described based on the LES solutions as well as providing a detailed analysis of the boundary layer behavior. The changes in shock structure, boundary layer interaction physics, and associated losses with operating condition are highlighted. A low-frequency SBLI unsteadiness is observed in the cases where the boundary layer into the shock is laminar, and a scaling of the frequency is proposed. The scaling is based on the time scale of turbulent structures convecting from the shock to the trailing edge and acoustic disturbances then traveling back upstream.
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Zhongwei, He, and Zhang Shiying. "Lip Separate Flow Blowing and Analysis of Coherence of Inlet." In ASME 1985 Beijing International Gas Turbine Symposium and Exposition. American Society of Mechanical Engineers, 1985. http://dx.doi.org/10.1115/85-igt-68.

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It is found in the experiments that blowing at the lip separation of the inlet obviously reduces the turbulences at the inlet exit, and apparently reduces the intensity of pressure fluctuations caused by the shock-boundary layer interaction down-stream of throat. The coherence between pressure in the interaction region and total pressure at the exit is also reduced. The coherence between the pressure in the lip separation region and total pressure at the exit is 0.32. If, in addition, there is a stronger shock down-stream of the throat the above mentioned coherence is reduced to 0.06.
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Aotsuka, Mizuho, Toshinori Watanabe, and Yasuo Machina. "Role of Shock and Boundary Layer Separation on Unsteady Aerodynamic Characteristics of Oscillating Transonic Cascade." In ASME Turbo Expo 2003, collocated with the 2003 International Joint Power Generation Conference. ASMEDC, 2003. http://dx.doi.org/10.1115/gt2003-38425.

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The unsteady aerodynamic characteristics of an oscillating compressor cascade composed of Double-Circular-Arc airfoil blades were both experimentally and numerically studied under transonic flow conditions. The study aimed at clarifying the role of shock waves and boundary layer separation due to the shock boundary layer interaction on the vibration characteristics of the blades. The measurement of the unsteady aerodynamic moment on the blades was conducted in a transonic linear cascade tunnel using an influence coefficient method. The cascade was composed of seven DCA blades, the central one of which was an oscillating blade in a pitching mode. The unsteady moment was measured on the central blade as well as the two neighboring blades. The behavior of the shock waves was visualized through a schlieren technique. A quasi-three dimensional Navier-Stokes code was developed for the present numerical simulation of the unsteady flow fields around the oscillating blades. A k-ε turbulence model was utilized to adequately simulate the flow separation phenomena caused by the shock-boundary layer interaction. The experimental and numerical results complemented each other and enabled a detailed understanding of the unsteady aerodynamic behavior of the cascade. It was found that the surface pressure fluctuations induced by the shock oscillation were the governing factor for the unsteady aerodynamic moment acting on the blades. Such pressure fluctuations were primarily induced by the movement of impingement point of the shock on the blade surface. During the shock oscillation the separated region caused by the shock boundary layer interaction also oscillated along the blade surface, and induced additional pressure fluctuations. The shock oscillation and the movement of the separated region were found to play the principal role in the unsteady aerodynamic and vibration characteristics of the transonic compressor cascade.
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Biswas, Debasish, and Tomohiko Jimbo. "Studies on Characteristic Frequency and Length Scale of Shock Induced Motion in Transonic Diffuser Using a High Order LES Approach." In ASME 2015 Gas Turbine India Conference. American Society of Mechanical Engineers, 2015. http://dx.doi.org/10.1115/gtindia2015-1225.

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Unsteady transonic flows in diffuser have become increasingly important, because of its application in new propulsion systems. In the development of supersonic inlet, air breathing propulsion systems of aircraft and missiles, detail investigations of these types of flow behavior are very much essential. In these propulsion systems, naturally present self-sustaining oscillations, believed to be equivalent to dynamically distorted flow fields in operational inlets, were found under all operating conditions. The investigations are also relevant to pressure oscillations known to occur in ramjet inlets in response to combustor instabilities. The unsteady aspects of these flows are important because the appearance of undesirable fluctuations generally impose limitation on the inlet performance. Test results of ramjet propulsion systems have shown undesirable high amplitude pressure fluctuations caused by the combustion instability. The pressure fluctuations originated from the combustor extend forward into the inlet and interact with the diffuser flow-field. Depending on different parameters such as the diffuser geometry, the inlet/exit pressure ratio, the flow Mach number, different complicated phenomena may occur. The most important characteristics are the occurrence of shock induced separation, the length of separation region downstream of the shock location, and the oscillation of shock location as well as the oscillation of the whole downstream flow. Sajben experimentally investigated in detail the time mean and unsteady flow characteristics of supercritical transonic diffuser as a function of flow Mach number upstream the shock location and diffuser length. The flows exhibited features similar to those in supersonic inlets of air-breathing propulsion systems of aircraft. A High-order LES turbulence model developed by the author is assessed with experimental data of Sajben on the self-excited shock oscillation phenomena. The whole diffuser model configuration including the suction slot located at certain axial location around the bottom and side walls to remove boundary layer, are included in the present computation model. The time-mean and unsteady flow characteristics in this transonic diffuser as a function of flow Mach number and diffuser length are investigated in detail. The results of study showed that in the case of shock-induced separation flow, the length and thickness of the reverse flow region of the separation-bubble change, as the shock passed through its cycle. The instabilities in the separated layer, the shock /boundary layer interaction, the dynamics of entrainment in the separation bubble, and the interaction of the travelling pressure wave with the pressure fluctuation region caused by the step-like structure of the suction slot play very important role in the shock-oscillation frequency.
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