Dissertations / Theses: 'Hypersonic boundary layer'

1

Williams, Simon. "Three-dimensional separation of a hypersonic boundary layer." Thesis, Imperial College London, 2005. http://hdl.handle.net/10044/1/11450.

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2

Wang, Xiaowen. "Numerical simulations of hypersonic boundary-layer stability and receptivity." Diss., Restricted to subscribing institutions, 2007. http://proquest.umi.com/pqdweb?did=1464122601&sid=1&Fmt=2&clientId=1564&RQT=309&VName=PQD.

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3

Manning, Melissa Lynn. "COMPUTATIONAL EVALUATION OF QUIET TUNNEL HYPERSONIC BOUNDARY LAYER STABILITY EXPERIMENTS." NCSU, 2001. http://www.lib.ncsu.edu/theses/available/etd-20010112-081130.

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Manning, Melissa Lynn. Computational Evaluation of Quiet Tunnel Hypersonic Boundary Layer Stability Experiments. (Under the direction of Dr. Ndaona Chokani.) A computational evaluation of two stability experiments conducted in the NASA Langley Mach 6 axisymmetric quiet nozzle test chamber facility is conducted. Navier-Stokes analysis of the mean flow and linear stability theory analysis of boundary layer disturbances is performed in the computations. The effects of adverse pressure gradient and wall cooling are examined. Calculated pressure, temperature and boundary layer thickness distributions show very good overall agreement with experimental measurements. Computed mass flux and total temperature profiles show very good quantitative agreement with uncalibrated hot-wire measurements obtained with the hot-wire operated in high and low overheat modes respectively. Comparisons between calibrated hot-wire data and mean flow computations show excellent agreement in the early stages of the transitional flow. However, examination of the wire Reynolds number and mass flux and total temperature eigenfunction profiles suggest that when operated in high overheat mode the sensitivity of the hot-wire to total temperature is significant. Thus, while uncalibrated hot-wire measurements are useful to characterize the overall features of the flow, calibrated hot-wire measurements are necessary for quantitative comparison with stability theory. Computations show that adverse pressure gradient and wall cooling decrease the boundary layer thickness and increase the frequency and amplification rate of the unstable second mode disturbances; these findings are consistent with the experimental observations.

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4

Surah, Davinder. "Investigation of attachment line boundary layer characteristics in hypersonic flows." Thesis, Cranfield University, 2000. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.323921.

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5

Atcliffe, Phillip Arthur. "Effects of boundary layer separation and transition at hypersonic speeds." Thesis, Cranfield University, 1995. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.336458.

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6

Tirtey, Sandy C. "Characterization of a transitional hypersonic boundary layer in wind tunnel and flight conditions." Doctoral thesis, Universite Libre de Bruxelles, 2009. http://hdl.handle.net/2013/ULB-DIPOT:oai:dipot.ulb.ac.be:2013/210367.

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Laminar turbulent transition is known for a long time as a critical phenomenon influencing the thermal load encountered by hypersonic vehicle during their planetary re-entry trajectory. Despite the efforts made by several research laboratories all over the world, the prediction of transition remains inaccurate, leading to oversized thermal protection system and dramatic limitations of hypersonic vehicles performances. One of the reasons explaining the difficulties encountered in predicting transition is the wide variety of parameters playing a role in the phenomenon. Among these parameters, surface roughness is known to play a major role and has been investigated in the present thesis.

A wide bibliographic review describing the main parameters affecting transition and their coupling is proposed. The most popular roughness-induced transition predictions correlations are presented, insisting on the lack of physics included in these methods and the difficulties encountered in performing ground hypersonic transition experiments representative of real flight characteristics. This bibliographic review shows the importance of a better understanding of the physical phenomenon and of a wider experimental database, including real flight data, for the development of accurate prediction methods.

Based on the above conclusions, a hypersonic experimental test campaign is realized for the characterization of the flow field structure in the vicinity and in the wake of 3D roughness elements. This fundamental flat plate study is associated with numerical simulations for supporting the interpretation of experimental results and thus a better understanding of transition physics. Finally, a model is proposed in agreement with the wind tunnel observations and the bibliographic survey.

The second principal axis of the present study is the development of a hypersonic in-flight roughness-induced transition experiment in the frame of the European EXPERT program. These flight data, together with various wind tunnel measurements are very important for the development of a wide experimental database supporting the elaboration of future transition prediction methods.
Doctorat en Sciences de l'ingénieur
info:eu-repo/semantics/nonPublished

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7

Bura, Romie Oktovianus. "Laminar/transitional shock-wave/boundary-layer interactions (SWBLIs) in hypersonic flows." Thesis, University of Southampton, 2004. https://eprints.soton.ac.uk/47605/.

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Numerical investigations of laminar shock-wave/boundary-layer interactions (SWBLIs) in hypersonic flow have been carried out at M∞ = 6.85 and M∞ ≈ 8, with unit Reynolds numbers ranging from 2.0 x 106 m- l to 7.60 x 106 m- l. This thesis deals with a simplified 2-D geometric configuration to simulate SWBLIs on vehicle surfaces or engine intakes, i.e. the interaction of an oblique shock (produced by a wedge) impinging on an incoming laminar boundary-layer on an isothermal flat plate. The numerical simulations were performed with weak/moderate to strong shock. The results were compared with available theoretical and experimental results. Limited experimental work at M∞ = 6.85 for obtaining qualitative data were performed to provide the location of separation and re-attachment points using surface oil flow. Schlieren photographs were taken to provide the general flow features. A comprehensive analysis was performed on the 2-D numerical results with various Mach numbers, Reynolds numbers and shock strengths, to verify whether numerical solutions were able to confirm the established trends for the laminar free-interaction concept. An analysis was also performed using a well-established power-law relationship of pressure and heat flux in the region of interactions. An unstable first oblique mode disturbance was imposed with the strongest wedge angle, 9°, at M∞ = 6.85 and unit Reynolds number 2.45 x 106 m- l to determine the boundary-layer stability and its propensity to undergo transition in the linear regime. Several unsteady 3-D simulations were performed with varied parameters. Streamwise vortices were generated in all cases especially downstream of maximum separation bubble height. However, as the amplifications of the disturbance were quite small, transition was found to be unlikely at these conditions

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8

Murray, Neil Paul. "Three-dimensional turbulent shock-wave : boundary-layer interactions in hypersonic flows." Thesis, Imperial College London, 2007. http://hdl.handle.net/10044/1/7963.

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9

Grossir, Guillaume. "Longshot hypersonic wind tunnel flow characterization and boundary layer stability investigations." Doctoral thesis, Universite Libre de Bruxelles, 2015. http://hdl.handle.net/2013/ULB-DIPOT:oai:dipot.ulb.ac.be:2013/209044.

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The hypersonic laminar to turbulent transition problem above Mach 10 is addressed experimentally in the short duration VKI Longshot gun tunnel. Reentry conditions are partially duplicated in terms of Mach and Reynolds numbers. Pure nitrogen is used as a test gas with flow enthalpies sufficiently low to avoid its dissociation, thus approaching a perfect gas behavior. The stabilizing effects of Mach number and nosetip bluntness on the development of natural boundary layer disturbances are evaluated over a 7 degrees half-angle conical geometry without angle of attack.

Emphasis is initially placed on the flow characterization of the Longshot wind tunnel where these experiments are performed. Free-stream static pressure diagnostics are implemented in order to complete existing stagnation point pressure and heat flux measurements on a hemispherical probe. An alternative method used to determine accurate free-stream flow conditions is then derived following a rigorous theoretical approach coupled to the VKI Mutation thermo-chemical library. Resulting sensitivities of free-stream quantities to the experimental inputs are determined and the corresponding uncertainties are quantified and discussed. The benefits of this different approach are underlined, revealing the severe weaknesses of traditional methods based on the measurement of reservoir conditions and the following assumptions of an isentropic and adiabatic flow through the nozzle. The operational map of the Longshot wind tunnel is redefined accordingly. The practical limits associated with the onset of nitrogen flow condensation under non-equilibrium conditions are also accounted for.

Boundary layer transition experiments are then performed in this environment with free-stream Mach numbers ranging between 10-12. Instrumentation along the 800mm long conical model includes flush-mounted thermocouples and fast-response pressure sensors. Transition locations on sharp cones compare favorably with engineering correlations. A strong stabilizing effect of nosetip bluntness is reported and no transition reversal regime is observed for Re_RN<120000. Wavelet analysis of wall pressure traces denote the presence of inviscid instabilities belonging to Mack's second mode. An excellent agreement with Linear Stability Theory results is obtained from which the N-factor of the Longshot wind tunnel in these conditions is inferred. A novel Schlieren technique using a short duration laser light source is developed, allowing for high-quality flow visualization of the boundary layer disturbances. Comparisons of these measurement techniques between each other are finally reported, providing a detailed view of the transition process above Mach 10.
Doctorat en Sciences de l'ingénieur
info:eu-repo/semantics/nonPublished

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10

Riley, Zachary Bryce Riley. "Interaction Between Aerothermally Compliant Structures and Boundary-Layer Transition in Hypersonic Flow." The Ohio State University, 2016. http://rave.ohiolink.edu/etdc/view?acc_num=osu1471618528.

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11

Husmeier, Frank. "Numerical Investigations of Transition in Hypersonic Flows over Circular Cones." Diss., The University of Arizona, 2008. http://hdl.handle.net/10150/196123.

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This thesis focuses on secondary instability mechanisms of high-speed boundary layers over cones with a circular cross section. Hypersonic transition investigations at Mach 8 are performed using Direct Numerical Simulations (DNS). At wind-tunnel conditions, these simulations allow for comparison with experimental measurements to verify fundamental stability characteristics.To better understand geometrical influences, flat-plate and cylindrical geometries are studied using after-shock conditions of the conical investigations. This allows for a direct comparison with the results of the sharp cone to evaluate the influence of the spanwise curvature and the cone opening angle. The ratio of the boundary-layer thickness to the spanwise radius is used to determine the importance of spanwise curvature effects. When advancing in the downstream direction the radius increaseslinearly while the boundary-layer thickness stays almost constant. Hence, spanwise curvature effects are strongest close to the nose and decrease in downstream direction. Their influences on the secondary instability mechanisms provide some rudimentary guidance in the design of future high-speed air vehicles.In experiments, blunting of the nose tip of the circular cone results in an increase in critical Reynolds number (c.f. Stetson et al. (1984)). However, once a certain threshold of the nose radius is exceeded, the critical Reynolds number decreases even to lower values than for the sharp cone. So far, conclusive explanations for this behavior could not be derived based on the available experimental data. Therefore, here DNS is used to study the effect of nose bluntness on secondary instability mechanisms in order to shed light on the underlying flow physics. To this end, three different nose tip radii are considered-the sharp cone, a small nose radius and a large nose radius. A small nose radius moves the transition on-set downstream, while for a large nose radius the so-called transition reversal is observed. Experimentalists hold influences of the entropy layer responsible but detailed numerical studies may lead to alternateconclusions.

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12

Aziz, Saduman. "Perfect Gas Navier-stokes Solutions Of Hypersonic Boundary Layer And Compression Corner Flows." Phd thesis, METU, 2005. http://etd.lib.metu.edu.tr/upload/12606661/index.pdf.

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The purpose of this thesis is to perform numerical solutions of hypersonic, high temperature, perfect gas flows over various geometries. Three dimensional, thin layer, compressible, Navier-Stokes equations are solved. An upwind finite difference approach with Lower Upper-Alternating Direction Implicit (LU-ADI) decomposition is used. Solutions of laminar, hypersonic, high temperature, perfect gas flows over flat plate and compression corners (qw=5°
, 10°
, 14°
, 15°
, 16°
, 18°
and 24°
) with eight different free-stream and wall conditions are presented and discussed. During the analysis, air viscosity is calculated from the Sutherland formula up to 1000°
K, for the temperature range between 1000 º
K and 5000 º
K a curve fit to the estimations of Svehla is applied. The effects of Tw/T0 on heat transfer rates, surface pressure distributions and boundary layer characteristics are studied. The effects of corner angle (&
#952
w) on strong shock wave/boundary layer interactions with extended separated regions are investigated. The obtained results are compared with the available experimental data, computational results, and theory.

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13

Lopez, Lee (Lee Gabriel). "Evaluation of a displacement-body model for hypersonic shock-wave/boundary-layer interaction." Thesis, Massachusetts Institute of Technology, 2018. http://hdl.handle.net/1721.1/119306.

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Thesis: S.M., Massachusetts Institute of Technology, Department of Aeronautics and Astronautics, 2018.
Cataloged from PDF version of thesis.
Includes bibliographical references (pages 127-128).
In this thesis, a displacement-body model for strong-interaction shock-wave/boundary-layer interaction (SBLI) is presented and evaluated. The model considers 2-D flow over a flat plate with an adiabatic wall. The separation bubble is modeled as a displacement body with constant surface pressure, the value of which is set equal to the value of plateau pressure given by free-interaction theory. A shock-fitting method of characteristics is employed to numerically compute quantities in the inviscid outer flow. Boundary conditions that satisfy physical requirements at shock waves, slip lines, and solid walls are enforced. Accuracy of the model is shown for both laminar and turbulent flow regimes, as well as for Mach numbers in the hypersonic regime. Additionally, the model provides a physical explanation for the pressure drop observed downstream of reattachment in hypersonic flows.
by Lee Lopez.
S.M.

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14

Cerminara, Adriano. "Boundary-layer receptivity and breakdown mechanisms for hypersonic flow over blunt leading-edge configurations." Thesis, University of Southampton, 2017. https://eprints.soton.ac.uk/412641/.

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Direct numerical simulations have been performed to study receptivity and breakdown mechanisms for hypersonic flow over blunt leading-edge configurations with imposed freestream acoustic disturbances. Both two-dimensional (2D) and three-dimensional (3D) fast and slow acoustic wave models have been used. The former has been adopted for 2D simulations over a blunt-nose wedge probe designed to measure freestream noise levels in hypersonic wind tunnels; the latter has been used to perform 3D simulations for a span-periodic blunt wedge in unswept and swept configurations, and for a three-dimensional generic forebody model. In the 2D wedge simulations, an analysis of the post-shock wave structure shows that fast acoustic waves are efficiently transmitted across the shock as refracted waves, while slow acoustic waves generate convected waves. The wall response to the fast mode highlights a resonance-modulation behaviour in the nose region. An estimation of the freestream noise levels in the DLR high-enthalpy (HEG) and low-enthalpy (RWG) hypersonic wind tunnels has been performed, showing higher noise levels for the HEG wind tunnel at high Mach numbers. The 3D wedge simulations have been used to study the characteristics of the receptivity and breakdown mechanisms associated with different wave types (fast/slow), disturbance amplitudes, and sweep angles. The fast-mode induced transition has been observed to be a much more rapid and powerful process than the slow-wave related transition, because of the role played by the fast-mode resonance mechanism at the leading edge. Finally, the numerical simulations performed for a generic forebody geometry have enabled comparison with a recent transition experiment carried out in the Mach 6 Purdue hypersonic wind tunnel in noisy conditions. In this case, slow acoustic waves show the most similar transition patterns to the experimental case, and, in particular, are more efficient than fast waves in triggering nonlinear growth of streamwise streaks, related to crossflow inflectional instabilities located in the off-centerline leading-edge region.

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15

Chuck, Chen. "Numerical simulation of oblique detonation and shock-deflagration waves with a laminar boundary-layer /." Thesis, Connect to this title online; UW restricted, 1990. http://hdl.handle.net/1773/9966.

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16

Mohammed, Sohail. "Experimental investigation of shock wave and boundary layer interaction near convex corners in hypersonic flow." Thesis, National Library of Canada = Bibliothèque nationale du Canada, 1997. http://www.collectionscanada.ca/obj/s4/f2/dsk2/ftp01/MQ28817.pdf.

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17

Schreyer, Anne-Marie [Verfasser]. "Experimental investigations of supersonic and hypersonic shock wave/turbulent boundary layer interactions / Anne-Marie Schreyer." München : Verlag Dr. Hut, 2013. http://d-nb.info/1045126853/34.

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18

Salemi, Leonardo da Costa, and Leonardo da Costa Salemi. "Numerical Investigation of Hypersonic Conical Boundary-Layer Stability Including High-Enthalpy and Three-Dimensional Effects." Diss., The University of Arizona, 2016. http://hdl.handle.net/10150/621854.

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The spatial stability of hypersonic conical boundary layers is investigated utilizing different numerical techniques. First, the development and verification of a Linearized Compressible Navier-Stokes solver (LinCS) is presented, followed by an investigation of different effects that affect the stability of the flow in free-flight/ground tests, such as: high-enthalpy effects, wall-temperature ratio, and three-dimensionality (i.e. angle-of-attack). A temporally/spatially high-order of accuracy parallelized Linearized Compressible Navier-Stokes solver in disturbance formulation was developed, verified and employed in stability investigations. Herein, the solver was applied and verified against LST, PSE and DNS, for different hypersonic boundary-layer flows over several geometries (e.g. flat plate - M=5.35 & 10; straight cone - M=5.32, 6 & 7.95; flared cone - M=6; straight cone at AoA = 6 deg - M=6). The stability of a high-enthalpy flow was investigated utilizing LST, LinCS and DNS of the experiments performed for a 5 deg sharp cone in the T5 tunnel at Caltech. The results from axisymmetric and 3D wave-packet investigations in the linear, weakly, and strongly nonlinear regimes using DNS are presented. High-order spectral analysis was employed in order to elucidate the presence of nonlinear couplings, and the fundamental breakdown of second mode waves was investigated using parametric studies. The three-dimensionality of the flow over the Purdue 7 deg sharp cone at M=6 and AoA =6 deg was also investigated. The development of the crossflow instability was investigated utilizing suction/blowing at the wall in the LinCS/DNS framework. Results show good agreement with previous computational investigations, and that the proper basic flow computation/formation of the vortices is very sensitive to grid resolution.

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19

Laible, Andreas Christian. "Numerical Investigation of Boundary-Layer Transition for Cones at Mach 3.5 and 6.0." Diss., The University of Arizona, 2011. http://hdl.handle.net/10150/205419.

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Transition in high-speed boundary layers is investigated using direct numerical simulation (DNS). A compressible Navier-Stokes code that is specifically tailored towards accurate and efficient simulations of boundary layer stability and boundary layer transition was developed and thoroughly validated. Particular emphasis was put into the adoption of a high-order accurate spatial discretization including a boundary closure with the same stencil width as the interior scheme. Oblique breakdown has been shown, using both temporal and spatial DNS, to be a viable route to transition for the boundary layer of the sharp 7° cone at Mach 3.5 investigated by Corke 2002. A 'wedge-shaped' transitional regime was observed to be characteristic for this type of breakdown on the cone geometry. Furthermore, it was shown that the dominance of the longitudinal mode in the nonlinear transition regime of oblique breakdown is due to a continuously nonlinear forced transient growth. That is the primary pair of oblique waves permanently 'seeds' disturbances into the longitudinal mode, where these disturbances exhibit non-modal unstable behavior. In addition to the simulations of controlled transition via oblique breakdown, six simulations have been conducted and analyzed where transition is initiated by multiple primary waves. Despite the broader spectrum of primary waves, typical features of oblique breakdown are still apparent in these simulations and therefore, it may be conjectured, that oblique breakdown initiated by one primary pair of waves is a good model for the nonlinear processes in natural transition. Furthermore, hypersonic boundary layer stability and transition for a flared and a straight cone at Mach 6 was investigated. In particular, a comparative investigation between both geometries regarding the K-type breakdown was performed in order to give some indications towards the open question how strong the nonlinear transition processis altered by the cone flare.

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20

Saad, Mohd Rashdan. "Experimental studies on shock boundary layer interactions using micro-ramps at Mach 5." Thesis, University of Manchester, 2013. https://www.research.manchester.ac.uk/portal/en/theses/experimental-studies-on-shock-boundary-layer-interactions-using-microramps-at-mach-5(71f1e11c-dbfd-443a-a9ee-e3fc160176f1).html.

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Shock boundary layer interactions (SBLI) is an undesirable event occurring in high-speed air-breathing propulsion system that stimulates boundary layer separation due to adverse pressure gradients and consequently lead to ow distortion and pressure loss in the intake section. Therefore it is essential to apply ow control mechanisms to prevent this phenomenon. This study involves a novel ow control device called micro-ramp, which is a part of the micro-vortex generator family that has shown great potential in solving the adverse phenomenon. The term micro refers to the height of the device, which is smaller than the boundary layer thickness, δ. It is important to highlight the two main novelties of this investigation. Firstly, it is the first micro-ramp study conducted in the hypersonic ow regime (Mach 5) since most of the previous micro-ramp studies were only performed in subsonic, transonic and supersonic flows. Another novelty is the various experimental techniques that were used in single study for example schlieren photography, oil-dot and oil- ow visualisation and conventional pressure transducers. In addition, advanced ow diagnostic tools such as infrared thermography, pressure sensitive paints (PSP) and particle image velocimetry (PIV) were also employed. T

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21

Sivasubramanian, Jayahar. "Numerical Investigation of Laminar-Turbulent Transition in a Cone Boundary Layer at Mach 6." Diss., The University of Arizona, 2012. http://hdl.handle.net/10150/228514.

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Direct Numerical Simulations (DNS) are performed to investigate laminar-turbulent transition in a boundary layer on a sharp cone at Mach 6. The main objective of this dissertation research is to explore which nonlinear breakdown mechanisms may be dominant in a broad--band "natural" disturbance environment and then use this knowledge to perform controlled transition simulations to investigate these mechanisms in great detail. Towards this end, a "natural" transition scenario was modeled and investigated by generating wave packet disturbances. The evolution of a three-dimensional wave packet in a boundary layer has typically been used as an idealized model for "natural" transition to turbulence, since it represents the impulse response of the boundary layer and, thus, includes the interactions between all frequencies and wave numbers. These wave packet simulations provided strong evidence for a possible presence of fundamental and subharmonic resonance mechanisms in the nonlinear transition regime. However, the fundamental resonance was much stronger than the subharmonic. In addition to these two resonance mechanisms, the wave packet simulations also indicated the possible presence of oblique breakdown mechanism. To gain more insight into the nonlinear mechanisms, controlled transition simulations were performed of these mechanisms. Several small and medium scale simulations were performed to scan the parameter space for fundamental and subharmonic resonance. These simulations confirmed the findings of the wave packet simulations, namely that, fundamental resonance is much stronger compared to the subharmonic resonance. Subsequently a set of highly resolved fundamental and oblique breakdown simulations were performed. In these DNS, remarkable streamwise arranged "hot'' streaks were observed for both fundamental and oblique breakdown. The streaks were a consequence of the large amplitude steady longitudinal vortex modes in the nonlinear régime. These simulations demonstrated that both second--mode fundamental breakdown and oblique breakdown may indeed be viable paths to complete breakdown to turbulence in hypersonic boundary layers at Mach 6.

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22

Yentsch, Robert J. "Three-Dimensional Shock-Boundary Layer Interactions in Simulations of HIFiRE-1 and HIFiRE-2." The Ohio State University, 2013. http://rave.ohiolink.edu/etdc/view?acc_num=osu1384195671.

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23

Padilla, Montero Ivan. "Analysis of the stability of a flat-plate high-speed boundary layer with discrete roughness." Doctoral thesis, Universite Libre de Bruxelles, 2021. https://dipot.ulb.ac.be/dspace/bitstream/2013/324490/5/contratPM.pdf.

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Boundary-layer transition from a laminar to a turbulent regime is a critical driver in the design of high-speed vehicles. The aerothermodynamic loads associated with transitional or fully turbulent hypersonic boundary layers are several times higher than those associated with laminar flow. The presence of isolated roughness elements on the surface of a body can accelerate the growth of incoming disturbances and introduce additional instability mechanisms in the flow field, eventually leading to a premature occurrence of transition. This dissertation studies the instabilities induced by three-dimensional discrete roughness elements located inside a high-speed boundary layer developing on a flat plate. Two-dimensional local linear stability theory (2D-LST) is employed to identify the instabilities evolving in the three-dimensional flow field that characterizes the wake induced by the roughness elements and to investigate their evolution downstream. A formulation of the disturbance energy evolution equation available for base flows depending on a single spatial direction is generalized for the first time to base flows featuring two inhomogeneous directions and perturbations depending on three spatial directions. This generalization allows to obtain a decomposition of the temporal growth rate of 2D-LST instabilities into the different contributions that lead to the production and dissipation of the total disturbance energy. This novel extension of the formulation provides an additional layer of information for understanding the energy exchange mechanisms between a three-dimensional base flow and the perturbations resulting from 2D-LST. Stability computations for a calorically perfect gas illustrate that the wake induced by the roughness elements supports the growth of different sinuous and varicose instabilities which coexist together with the Mack-mode perturbations that evolve in the flat-plate boundary layer, and which become modulated by the roughness-element wake. A single pair of sinuous and varicose disturbances is found to dominate the wake instability in the vicinity of the obstacles. The application of the newly developed decomposition of the temporal growth rate reveals that the roughness-induced wake modes extract most of their potential energy from the transport of entropy fluctuations across the base-flow temperature gradients and most of their kinetic energy from the work of the disturbance Reynolds stresses against the base-flow velocity gradients. Further downstream, the growth rate of the wake instabilities is found to be influenced by the presence of Mack-mode disturbances developing on the flat plate. Strong evidence is observed of a continuous synchronization mechanism between the wake instabilities and the Mack-mode perturbations. This phenomenon leads to an enhancement of the amplification rate of the wake modes far downstream of the roughness element, ultimately increasing the associated integrated amplification factors for some of the investigated conditions. The effects of vibrational molecular excitation and chemical non-equilibrium on the instabilities induced by a roughness element are studied for the case of a high-temperature boundary layer developing on a sharp wedge configuration. For this purpose, a 2D-LST solver for chemical non-equilibrium flows is developed for the first time, featuring a fully consistent implementation of the thermal and transport models employed for the base flow and the perturbation fields. This is achieved thanks to the automatic derivation and implementation tool (ADIT) available within the von Karman Institute extensible stability and transition analysis (VESTA) tool-kit, which enables an automatic derivation and implementation of the 2D-LST governing equations for different thermodynamic flow assumptions and models. The stability computations for this configuration show that sinuous and varicose disturbances also dominate the wake instability in the presence of vibrational molecular energy mode excitation and chemical reactions. The resulting base-flow cooling associated with the modeling of such high-temperature phenomena is found to have opposite stabilizing and destabilizing effects on the streamwise evolution of the sinuous and varicose instabilities. The modeling of vibrational excitation and chemical non-equilibrium acting exclusively on the perturbations is found to have a stabilizing influence in all cases.
Doctorat en Sciences de l'ingénieur et technologie
info:eu-repo/semantics/nonPublished

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24

北村, 圭一, Keiichi KITAMURA, 佳朗中村, and Yoshiaki NAKAMURA. "極超音速衝撃波干渉流れにおける空力加熱の数値解析." 日本航空宇宙学会, 2008. http://hdl.handle.net/2237/13871.

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北村, 圭一, Keiichi KITAMURA, 啓伺小澤, Hiroshi OZAWA, 勝祥花井, Katsuhisa HANAI, 浩一森, Koichi MORI, 佳朗中村, and Yoshiaki NAKAMURA. "極超音速TSTOにおける衝撃波干渉・境界層剥離を伴う流れ場の解析." 日本航空宇宙学会, 2008. http://hdl.handle.net/2237/13872.

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26

Di, Giovanni Antonio [Verfasser], Christian [Akademischer Betreuer] Stemmer, Wolfgang [Gutachter] Schröder, and Christian [Gutachter] Stemmer. "Roughness-Induced Transition in a Hypersonic Capsule Boundary Layer under Wind-Tunnel and Reentry Conditions / Antonio Di Giovanni ; Gutachter: Wolfgang Schröder, Christian Stemmer ; Betreuer: Christian Stemmer." München : Universitätsbibliothek der TU München, 2020. http://d-nb.info/1211725227/34.

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27

Brouwer, Kirk Rowse. "Enhancement of CFD Surrogate Approaches for Thermo-Structural Response Prediction in High-Speed Flows." The Ohio State University, 2018. http://rave.ohiolink.edu/etdc/view?acc_num=osu1543340520905498.

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28

西野, 敦洋, Atsuhiro NISHINO, 尊史石川, Takahumi ISHIKAWA, 圭一北村, Keiichi KITAMURA, 佳朗中村, and Yoshiaki NAKAMURA. "極超音速TSTO空力干渉流れ場における2物体間隔の空力加熱率への影響." 日本航空宇宙学会, 2005. http://hdl.handle.net/2237/13879.

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29

Vanstone, Leon. "Shock-induced separation of transitional hypersonic boundary layers." Thesis, Imperial College London, 2014. http://hdl.handle.net/10044/1/24803.

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This thesis presents a joint experimental/CFD investigation of shock-induced boundary layer separations in hypersonic transitional boundary layers with an emphasis on collapse and re-establishment times of the separation bubble. This study also provides high fidelity measurements and excellent characterisation of the flow field in order to provide benchmark data of a challenging flow configuration with which to benchmark next generation CFD solvers. The experiments were conducted in the Imperial College Aeronautics Department Number Two Gun Tunnel, a Mach 8.9 axisymmetric facility with a freestream unit Reynolds number of 47 million An axisymmetric blunt-nosed cylinder fitted with an 8 degree flare forms the primary vehicle for this study, although a 1.3 degree cowl geometry was also used to impinge a shock onto the blunt-nosed cylinder. The shock boundary layer interaction was designed such that it was separated for a laminar boundary layer and collapsed for a turbulent one. Carefully controlled turbulent spots were generated upstream of the interaction region which passed through the separation causing its collapse and subsequent re-establishment. Two intermittency cases are considered, one where turbulent spot spacing is large and collapse/re-establishment pairs can be considered independent of each other and one where they can not. Experimental surface quantities through the interaction region are measured using either heat-transfer or pressure measurements and schlieren video is used to diagnose the larger shock structure. Further a non-intrusive toluene PLIF method is assessed for use in this facility and shows promise. CFD simulations are done using an in-house operator split Godunov solver with a Baldwin-Lomax turbulence model. CFD simulations show good agreement with experiment and provides information on flow quantities that would be extremely difficult to measure otherwise. Collapse times of the separation bubble were found to be fast in relation to characteristic spot passage times. The collapse process is also fast in relation to the surrounding flows ability to adjust, with collapse associated with significant shock curvature of the immediate outboard shock structures. This leads to unsteadiness, with surface pressure measurements exceeding the range bounded by the laminar separated and turbulent collapsed cases. The severity of the unsteadiness appears to be driven by turbulent spot spacing. Re-establishment is considerably slower, showing asymptotic recovery that is likely driven by viscous diffusion rates, taking many characteristic spot passage times to recover.

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Babinsky, Holger. "A study of roughness in turbulent hypersonic boundary-layers." Thesis, Cranfield University, 1993. http://dspace.lib.cranfield.ac.uk/handle/1826/7586.

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The influence of large scale regular roughness on a Mach 5 turbulent boundary layer and a compression corner was investigated on axisymmetric wind tunnel models. Three types of roughness were examined; a series of square cavities at two different sizes and a 45 degree sawtooth. Typical sizes ranged from 50% to 100% of an undisturbed boundary layer thickness. The roughness was limited to a short region followed by a smooth surface. Compression corners were formed by 15° and 20° flares located downstream of the roughness. The flow in the wind tunnel was investigated in detail to obtain knowledge on operating conditions and flow quality. Liquid crystal thermography was developed for routine use in hypersonic blow-down wind tunnels with superior spatial resolution and experimental uncertainties in the range of traditional techniques. The effect on flow parameters downstream of the last roughness element were 7, found to differ significantly for the different quantities. Velocity profiles were found i, to be less full and skin friction was found to be reduced for all streamwise "~ distances. Surface heat transfer was increased in a short region limited to 1.5 boundary layer thicknesses behind the roughness whereas surface pressure was not affected. Sawtooth shaped roughness was found to cause a stronger j disturbance than square cavities of twice the size. Little influence of the roughness was noted on the flow over the compression corner. The flow over the 20° compression corner showed an increase in upstream influence for the sawtooth shaped roughness as well as the larger cavities. Surface pressure measurements did not indicate a separation in any case. Heat transfer measurements revealed a peak located approximately 0.25 boundary layer thicknesses behind the corner. No such feature was found in the surface pressure distributions. It is suggested that a small scale separation is located very close to the corner causing the peak in heat transfer at reattachment without any effect on surface pressures. The existence of such a separation has been confirmed by surface flow visualisations for both flares.

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DiGregorio, Nicholas J. "Characteristics of Turbulent Boundary Layers along a Hypersonic Vehicle." Thesis, State University of New York at Buffalo, 2018. http://pqdtopen.proquest.com/#viewpdf?dispub=10822170.

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The flight conditions of a hypersonic vehicle on an ascent trajectory are computed and Reynolds-averaged Navier-Stokes (RANS) simulations of the turbulent boundary layers are performed across a Mach number range of 0.3 up to 16 using the computational fluid dynamics (CFD) software, VULCAN. The boundary conditions and leading edge geometry are varied from the simple case of adiabatic and sharp to cooled and blunted to reveal the physics of how these effects impact the results of flat plate boundary layer methods as applied to practical aerospace systems. The law of the wall, the Van Driest transformation, and a shear stress preserving transformation's ability to collapse boundary layer velocity profiles under the conditions of variable wall boundary condition and leading edge geometry is explored.

Boundary layer related quantities examined include the boundary layer thickness, local skin friction coefficient, displacement thickness, momentum thickness, heat flux, and integrated loads. It is found that cooling the surface serves to increase the density of the boundary layer, making it thinner. This thinning of the boundary layer thickness increases the velocity gradients, thus increasing the shear stresses and the local skin friction coefficient. The effects on turbulent boundary layers of blunting the leading edge are explained by the difference in properties, particularly viscosity, caused by a detached bow shock instead of a Mach wave that comes off of a sharp nose plate. Heat flux into a vehicle is found to be insignificant at low speeds, but increases drastically as the Mach number rises into the supersonic and hypersonic regimes. It is observed that the integrated skin friction coefficient decreases as Mach number increases and the leading edge becomes blunted, however, it increases as more cooling is applied at the boundary. The integrated heat flux computed from a sharp leading edge geometry is greater compared to a blunted leading edge due to greater temperature gradients in the sharp nose case relative to the blunt nose case.

The shear stress preserving transformation, derived with the inclusion of a stress balance condition, is found to produce a better collapse of the velocity profile data than the Van Driest transformation and the incompressible law of the wall regardless of Mach number, boundary condition or leading edge geometry. The normalized untransformed velocity gradients are compared to the velocity gradients resulting from the Van Driest and shear stress preserving tranformation. It is shown that the velocity gradients from the shear stress preserving match the normalized untransformed velocity gradients more closely than the Van Driest velocity gradients do. The advantages, disadvantages, and limitations of each transformation are discussed.

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Denman, Paul Ashley. "Experimental study of hypersonic boundary layers and base flows." Thesis, Imperial College London, 1996. http://hdl.handle.net/10044/1/45466.

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This experimental study documents the development and separation of a hypersonic boundary layer produced naturally on the cold surface of a sharp slender cone. At the base of the conical forebody, the equilibrium turbulent boundary layer was allowed to separate over an axisymmetric rearward facing step to form a compressible base flow. The investigation was conducted in the Imperial College No.2 gun tunnel at a freestream Mach number of 9 and unit Reynolds numbers of 15 and 55 million. The compressible boundary layer study was carried out at both of the available freestream unit Reynolds numbers and the measured data include distributions of wall static pressure and heat transfer rate, together with profiles of pitot pressure through the boundary layer. Using the chordwise distribution of surface heat flux as a means of transition detection, the cone transition Reynolds number was found to be 5.4x10^. This result, together with that obtained from flat plate studies conducted in the same test facility, provided a ratio of cone to flat plate transition Reynolds number of 0.8. Boundary layer integral quantities and shape factors are derived from velocity profiles and in most cases the measured data extended close enough to the wall to detect the peak values of the integrands. The separated flow region formed at the base of the cone was documented only at the higher unit Reynolds number, a condition under which the approaching turbulent boundary layer was found to be close to equilibrium. The data include pitot pressure profiles recorded normal to the surface downstream of reattachment, together with wall static pressure and heat transfer rate distributions measured throughout the base flow region. Reattachment occurred approximately two step heights downstream of separation and a surface flow visualisation study indicated the existence of Taylor-Goertler type vortices, emanating from the reattachment line in the downstream direction. A simple shear layer expansion model is developed and shown to provide a favourable prediction of the measured pitot pressure profiles recorded downstream of the reattachment line. The success of this second order model implies that the dynamics of the corner expansion process, except in the immediate vicinity of the wall, is governed largely by inviscid pressure mechanisms and that the supersonic region of the boundary layer expansion is essentially isentropic.

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Van, den Eynde Jeroen. "Stability and transition of the flow behind isolated roughness elements in hypersonic boundary layers." Thesis, University of Southampton, 2015. https://eprints.soton.ac.uk/386204/.

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In this work the effect of isolated surface roughness on the behaviour of a hypersonic boundary layer is investigated, with a particular focus on the effect of the three-dimensional roughness shape on the instability of the roughness wake and the subsequent transition process. The analysis is performed computationally using direct numerical simulations, which solve the compressible Navier-Stokes equations, and a new code, developed in the scope of the current work, to analyse the linear stability of these equations. The full three-stage roughness-induced transition process has been investigated: firstly, the receptivity process and generation of boundary layer instabilities from freestream disturbances; secondly, the generation of a roughness wake and its initial linear instability; and finally the non-linear breakdown to turbulence of the roughness wake. In particular the effect of the three-dimensional roughness shape on these processes has been studied, looking at the roughness height, frontal profile, planform shape and upward/downward ramps. Also the effect of freestream disturbance amplitude andwall cooling has been investigated. It has been found that the roughness height and frontal profile have a large influence on the stability characteristics of the resulting wake and the subsequent transition. The roughness planform shape has a marginal effect, although cylindrical and diamond-shaped elements yield more unstable wakes than a square roughness element. Bi-local stability analysis can be used in most cases to predict the most unstable wake mode, but it under-predicts the instability growth rates due to non-parallel effects. The roughness shape has been observed to affect the transition onset location. The criteria commonly used to predict roughness-induced transition, do not take into account the three-dimensional shape, and an alternative transition prediction, based on the amplitude of the roughness-induced streamwise streak, is considered.

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Buguin, Arnaud. "Couches limites tridimensionnelles en hypersonique : effets du déséquilibre et du gradient d'entropie." Toulouse, ENSAE, 1997. http://www.theses.fr/1997ESAE0003.

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Les vitesses atteintes par les corps de rentrée sont très élevées, l'énergie cinétique est considérable. La quantité de chaleur mise en jeu est telle que l'écoulement hyperenthalpique devient le siège de processus physico-chimiques. Pour représenter l'écoulement, il est indispensable de préciser les interactions entre les phénomènes de déséquilibre thermochimique et les conditions aérodynamiques qui se traduisent par le couplage des équations cinétiques et de conservation. Les résultats de la thermodynamique statistique sont utilisés pour décrire les caractéristiques macroscopiques des espèces gazeuses. Les coefficients de transport sont obtenus selon la théorie cinétique de Boltzmann par des formulations simplifiées mais aussi plus générales pour le cas de mélanges complexes. La détermination des termes de production chimique et les différentes contributions aux termes sources des équations de relaxation vibrationnelle sont présentées en s'appuyant sur le schéma cinétique de l'air. La résolution des équations de couche limite de Prandtl ne permettant pas de tenir compte du rotationnel de l'écoulement extérieur, l'approche déficitaire a été étendue au cas des écoulements en déséquilibre. Cette technique repose sur la méthode des développements asymptotiques raccordés et permet d'assurer le raccord entre la solution visqueuse et la zone externe de fluide parfait. La mise en oeuvre numérique repose sur la méthode originale du code 3C3D permettant de conserver la forme cartésienne des équations de couche limite. Cette technique a été adaptée pour prendre en compte les phénomènes de relaxation et la formulation déficitaire. Enfin, des résultats de validation et d'application sont présentés sur différentes configurations.

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Chpoun, Amer. "Contribution a l'etude d'ecoulements hypersoniques (m=5) sur une rampe de compression en configuration 2-d et 3-d." Paris 6, 1988. http://www.theses.fr/1988PA06A005.

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Etude experimentale de l'influence de l'ecoulement transversal sur les distributions du flux thermique et de la pression parietale. Determination des grandeurs caracteristiques de la zone de decollement. Etude de l'apparition de la transition dans la zone du decollement en fonction du nombre de reynolds. Solution numerique pour la distribution de pression dans le cas de l'interaction laminaire

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36

Michael, Vipin George. "Effects of passive porous walls on the first Mack mode instability of hypersonic boundary layers over a sharp cone." Thesis, University of Birmingham, 2012. http://etheses.bham.ac.uk//id/eprint/3750/.

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Passive porous coatings have been proposed in literature as a means of delaying transition to turbulence in hypersonic boundary layers. The nonlinear stability of hypersonic viscous flow over a sharp slender cone with passive porous walls is investigated in this study. Hypersonic flows are unstable to viscous and inviscid disturbances, and following Mack (1984) these have been called the first and second Mack modes. A weakly nonlinear analysis of the instability of the flow to axisymmetric and non-axisymmetric viscous (first Mack mode) disturbances is performed here. The attached shock and effect of curvature are taken into account. Asymptotic methods are used at large Reynolds number and large Mach number to examine the viscous modes of instability, which may be described by a triple-deck structure. Various porous wall models have been incorporated into the stability analysis. The eigenrelations governing the linear stability of the problem are derived. Neutral and spatial instability results show the presence of multiple unstable modes and the destabilising effect of the porous wall models on them. The weakly nonlinear stability analysis carried out allows an equation for the amplitude of disturbances to be derived. The stabilising or destabilising effect of nonlinearity is found to depend on the cone radius. It is shown that porous walls significantly influences the effect of nonlinearity. They allow nonlinear effects to destabilise linearly unstable lower frequency modes and stabilise linearly unstable higher frequency modes.

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37

Sussman, Darien. "The influence of equivalence ratio and wall temperature on the ignition of H¦2/air mixtures in hypersonic flow boundary layers." Thesis, National Library of Canada = Bibliothèque nationale du Canada, 1998. http://www.collectionscanada.ca/obj/s4/f2/dsk1/tape11/PQDD_0009/MQ40900.pdf.

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38

Chpoun, Amer. "Contribution à l'étude d'écoulements hypersoniques (M=5) sur une rampe de compression en configuration 2-D et 3-D." Paris 6, 1988. http://www.theses.fr/1988PA066149.

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Etude expérimentale de l'influence de l'écoulement transversal sur les distributions du flux thermique et de la pression pariétale. Détermination des grandeurs caractéristiques de la zone de décollement. Etude de l'apparition de la transition dans la zone du décollement en fonction du nombre de Reynolds. Solution numérique pour la distribution de pression dans le cas de l'interaction laminaire

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André, Thierry. "Contrôle actif de la transition laminaire-turbulent en écoulement hypersonique." Thesis, Orléans, 2016. http://www.theses.fr/2016ORLE2022/document.

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Lors d’un vol hypersonique (Mach 6, 20 km d’altitude) la couche limite se développant sur l’avant-corps d’un véhicule hypersonique est laminaire. Cet état cause un désamorçage du moteur (statoréacteur) assurant la propulsion du véhicule. Pour pallier ce problème, il faut forcer la transition de la couche limite á l’aide d’un dispositif de contrôle dont l’effet est permanent (passif) ou modulable (actif) pendant le vol. Dans ce travail, nous analysons l’efficacité d’un dispositif actif d’injection d’air á la paroi pour forcer la transition de la couche limite sur un avant-corps générique. L’interaction jet d’air/couche limite est simulée numériquement avec une approche aux grandes échelles (LES). Une étude paramétrique sur la pression d’injection permet de quantifier l’efficacité du jet á déstabiliser la couche limite. L’influence des conditions de vol (altitude, Mach) sur la transition est également étudiée. Une analyse des résultats de simulation par Décomposition en Modes Dynamiques (DMD) est menée pour comprendre quels sont les modes dynamiques responsables de la transition et les mécanismes sous-jacents. Des essais dans la soufflerie silencieuse de l’université de Purdue (BAM6QT) ont été effectués pour tester expérimentalement l’efficacité des dispositifs passifs (rugosité isolée en forme de losange) et actifs (mono-injection d’air) pour faire transitionner la couche limite. Une peinture thermo-sensible et des capteurs de pression (PCB, Kulite) ont été utilisés pour déterminer la nature de la couche limite. Les résultats de ce travail montrent qu’une injection sonique suffit pour forcer la couche limite. On observe des essais, que pour une même hauteur de pénétration, les rugosités isolées sont moins efficaces que les jets (mono injection) pour déstabiliser la couche limite
During a hypersonic flight (Mach 6, 20 km altitude), the boundary layer developing on the forebody of a vehicle is laminar. This state may destabilize the scramjet engine propelling the vehicle. To overcome this problem during the flight, the boundary layer transition has to be forced using a control device whose effect is fixed (passive) or adjustable (active). In this work, we analyze the efficiency of a jet in crossflow in forcing the boundary layer transition on a generic forebody. The flow is computed with a Large Eddy Simulations (LES) approach. A parametric study of the injection pressure allows the efficiency of the jet in tripping the boundary layer to be quantified. The influence of flight conditions (Mach, altitude) on the transition is also studied. Dynamic Mode Decomposition (DMD) is applied to the simulation results to determine the transition leading to dynamic modes and to understand underlying transition mechanisms. Experiments in the Purdue University quiet wind tunnel (BAM6QT) were performed to quantify the efficiency of a passive transition device (diamond roughnesses) and an active transition device (single air jet) in tripping the boundary layer. A thermo-sensitive paint and pressure transducers (Kulite, PCB) were used to determine the state of the boundary layer on the generic forebody. Experimental and numerical results show a sonic injection is sufficient to induce transition. We observe from the experiments that for the same penetration height, a single roughness is less efficient than a single air jet in destabilizing the boundary layer

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(8793053), Gregory R. McKiernan. "Instability and Transition on a Sliced Cone with a Finite-Span Compression Ramp at Mach 6." Thesis, 2020.

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Initial experiments on separated shock/boundary-layer interactions were carried out within the Boeing/AFOSR Mach-6 Quiet Tunnel. Measurements were made of hypersonic laminar-turbulent transition within the separation above a compression corner. This wind tunnel features freestream fluctuations that are similar to those in

flight. The present work focuses on the role of traveling instabilities within the shear layer above the separation bubble.

A 7 degree half-angle cone with a slice and a finite-span compression ramp was designed and tested. Due to a lack of space for post-reattachment sensors, early designs of this

generic geometry did not allow for measurement of a post-reattachment boundary layer. Oil flow and heat transfer measurements showed that by lengthening the ramp, the post-reattachment boundary layer could be measured. A parametric study was completed to determine that a 20 degree ramp angle caused reattachment at 45% of the

total ramp length and provided the best flow field for boundary-layer transition measurements.

Surface pressure fluctuation measurements showed post-reattachment wave packets and turbulent spots. The presence of wave packets suggests that a shear-layer

instability might be present. Pressure fluctuation magnitudes showed a consistent transition Reynolds numbers of 900000, based on freestream conditions and distance

from the nosetip. Pressure fluctuations grew exponentially from less than 1% to roughly 10% of tangent-wedge surface pressure during transition.

A high-voltage pulsed plasma perturber was used to introduce controlled disturbances into the boundary layer. The concept was demonstrated on a straight 7 degree half-angle circular cone. The perturbations successfully excited the second-mode instability at naturally unstable frequencies. The maximum second-mode amplitudes prior to transition were measured to be about 10% of the mean surface static pressure.

The plasma perturber was then used to disturb the boundary layer just upstream of the separation bubble on the cone with the slice and ramp. A traveling instability was measured post-reattachment but the transition location did not change for any tested condition. It appears that the excited shear-layer instability was not the dominant mechanism of transition.

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Lo, Wei-Jen, and 羅偉仁. "Simulation of Hypersonic Shock-Boundary Layer Interaction." Thesis, 2011. http://ndltd.ncl.edu.tw/handle/65104448648249790865.

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碩士
逢甲大學
航太與系統工程所
99
This paper describes using finite volume method to solve Navier-Stokes Equations about a hypersonic intake has a double-ramp compress section. Considering varied computing methods to process 2-D computing simulations of flow field. Computing simulations of the intake wall temperature rise and second intake face angle rise, we have further research about the flow field change result from the shock wave/boundary layer interaction. During the research, we test the computing flux type first, the final details show that the computing outcomes of Roe-FDS and AUSM flux type are similar, since the Roe-FDS has more stable computing process than AUSM, we choose Roe-FDS to be our computing flux type. About mesh test, we increase the mesh number at the leading edge of the intake, corner between two faces and wall. We also try to use mesh adaptation method to raise mesh number by the density contour surface gradient. About turbulence computing model test, simulation outcomes show only the SST k-ω computing model with the low Reynolds number correction could simulate the physics phenomenon of the shock wave/boundary layer interaction for the hypersonic air flow through the corner of two ramps in intake. The outcomes of the computing for the increasing intake wall temperature show that the range of the shock wave/boundary layer interaction region becomes bigger related with the increasing wall temperature, it has the biggest change when temperature rises from 300K to 600K. We also considering the effect of the variable specific heat ratio to the flow field characteristics in the increasing wall temperature computing, the effect fewer when the wall temperature is lower, contrariwise, the effect obvious when the wall temperature higher. We also compare the outcomes of increasing intake second ramp angle and increasing wall temperature, the range of the shock wave/boundary layer interaction region is smaller when the second face angle increases test case. The Stanton number pattern of the intake is mainly affected by the shock wave/boundary layer interaction, when the interaction between shock wave and boundary transits from laminar boundary to the turbulence boundary; it clearly shows the boundary layer separation forward with the increasing angle.

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English, Benjamin L. "Large-Scale Streamwise Turbulent Structures in Hypersonic Boundary Layers." Thesis, 2013. http://hdl.handle.net/1969.1/149550.

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Prior research in the field of boundary layer turbulence has identified streamwise-elongated large-scale turbulence structures in both low speed compressible and high speed (M=2.0) flow. No experimental work has been done in any flow of M> or =3 in an attempt to identify the presence or quantify the behavior of these structures, nor has any study of favorable pressure gradient or surface roughness element effects on these structures been conducted. This research used high-resolution Particle Imaging Velocimetry in a M = 4.9 blow-down wind tunnel accompanied by a series of data analysis in order to identify the existence of streamwise-elongated large-scale turbulence structures in a hypersonic boundary layer. Furthermore, this study identified physical and statistical behavior which suggests that increasing favorable pressure gradient had a substantial impact on both the structural coherence and relative intensity of these turbulent structures at all boundary layer heights tested. This experiment also identified similar effects on these structures in the lower half of the boundary layer as a result of the introduction of surface roughness elements. Finally, several trends were identified between the averaged turbulence statistics and the behavior of the large-scale streamwise-elongated turbulence structures present in this study.

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Zhikharev, Constantin N. "Interaction theory for hypersonic separation and supersonic flow past a flexible wall /." Diss., 1999. http://gateway.proquest.com/openurl?url_ver=Z39.88-2004&rft_val_fmt=info:ofi/fmt:kev:mtx:dissertation&res_dat=xri:pqdiss&rft_dat=xri:pqdiss:9935191.

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(6624017), Joshua B. Edelman. "Nonlinear Growth and Breakdown of the Hypersonic Crossflow Instability." Thesis, 2019.

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A sharp, circular 7° half-angle cone was tested in the Boeing/AFOSR Mach-6 Quiet Tunnel

at 6° angle of attack, extending several previous experiments on the growth and breakdown of

stationary crossflow instabilities in the boundary layer.

Measurements were made using infrared

imaging and surface pressure sensors. Detailed measurements of the stationary and traveling

crossflow vortices, as well as various secondary instability modes, were collected over a large

region of the cone.

The Rod Insertion Method (RIM) roughness, first developed for use on a flared cone, was

adapted for application to crossflow work. It was demonstrated that the roughness elements were

the primary factor responsible for the appearance of the specific pattern of stationary streaks

downstream, which are the footprints of the stationary crossflow vortices. In addition, a roughness

insert was created with a high RMS level of normally-distributed roughness to excite the naturally

most-amplified stationary mode.

The nonlinear breakdown mechanism induced by each type of roughness appears to be

different. When using the discrete RIM roughness, the dominant mechanism seems to be the

modulated second mode, which is significantly destabilized by the large stationary vortices. This

is consistent with recent computations. There is no evidence of the presence of traveling crossflow

when using the RIM roughness, though surface measurements cannot provide a complete picture.

The modulated second mode shows strong nonlinearity and harmonic development just prior

to breakdown. In addition, pairs of hot streaks merge together within a constant azimuthal

band, leading to a peak in the heating simultaneously with the peak amplitude of the measured

secondary instability. The heating then decays before rising again to turbulent levels. This nonmonotonic

heating pattern is reminiscent of experiments on a flared cone and earlier computations

of crossflow on an elliptic cone.

When using the distributed roughness there are several differences in the nonlinear breakdown

behavior. The hot streaks appear to be much more uniform and form at a higher wavenumber,

which is expected given computational results. Furthermore, the traveling crossflow waves become

very prominent in the surface pressure fluctuations and weakly nonlinear. In addition there

appears in the spectra a higher-frequency peak which is hypothesized to be a type-I secondary instability

under the upwelling of the stationary vortices. The traveling crossflow and the secondary

instability interact nonlinearly prior to breakdown.

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45

Semper, Michael Thomas. "Examining A Hypersonic Turbulent Boundary Layer at Low Reynolds Number." Thesis, 2013. http://hdl.handle.net/1969.1/150983.

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The purpose of the current study was to answer several questions related to hypersonic, low Reynolds number, turbulent boundary layers, of which available data related to turbulence quantities is scarce. To that end, a unique research facility was created, instrumentation was developed to acquire data in the challenging low Reynolds number (low density) domain, and meaningful data was collected and analyzed. The low Reynolds number nature of the boundary layer (Re_theta = 3700) allows for tangible DNS computations/validations using the current geometry and conditions. The boundary layer examined in this experiment resembled other, higher Reynolds number boundary layers, but also exhibited its own unique characteristics. The Van Driest equivalent velocity scaling method was found to perform well, and the log layer of the law of the wall plot matched expected theory. Noticeably absent from the data was an overlap region between the two layers, which suggests a different profile for the velocity profiles at these low Reynolds number, hypersonic conditions. The low density effects near the wall may be having an effect on the turbulence that modifies this region in a manner not currently anticipated. The Crocco-Busemann relation was found to provide satisfactory results under its general assumptions. When compared to available data, the Morkovin scaled velocity fluctuations fell almost an order of magnitude short. Currently, it is not known if this deficit is due to inadequacies with the Strong Reynolds Analogy, or the Morkovin scaling parameters. The trips seem to promote uniformity across the span of the model, and the data seems to generally be in agreement across the spanwise stations. However, additional information is needed to determine if two-dimensional simulations are sufficient for these boundary layers. When the turbulent boundary layer power spectra is analyzed, the result is found to follow the traditional power law. This result verifies that even at low Reynolds numbers, the length scales still follow the behavior described by Kolmogorov. Moving downstream of the trips, the peak RMS disturbance value grows in amplitude until it reaches a critical value. After this point, the peak begins to decrease in amplitude, but the affected region spreads throughout the boundary layer. Once the influenced region covers a significant portion of the boundary layer, transition occurs.

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Sriram, R. "Shock Tunnel Investigations on Hypersonic Impinging Shock Wave Boundary Layer Interaction." Thesis, 2013. http://etd.iisc.ernet.in/2005/3401.

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The interaction of a shock wave and boundary layer often occurs in high speed flows. For sufficiently strong shock strengths the boundary layer separates, generating shock patterns in the contiguous inviscid flow (termed strong interactions); which may also affect the performances of the systems where they occur, demanding control of the interaction to enhance the performances. The case of impinging shock wave boundary layer interaction is of fundamental importance and can throw light on the physics of the interaction in general. Although various aspects of the interaction are studied at supersonic speeds, the complexities involved in the interaction at hypersonic speeds are not well understood. Of importance is the high total enthalpy associated with hypersonic flows the simulation of which requires shock tunnels. The present experimental study focuses on the interaction between strong impinging shock and boundary layer in hypersonic flows of moderate to high total enthalpies. Experiments are performed in hypersonic shock tunnels HST-2 and FPST (free piston driven shock tunnel), at nominal Mach numbers 6 and 8, with total enthalpy ranging from 1.3 MJ/kg to 6 MJ/kg, and freestream Reynolds number ranging from 0.3 million/m to 4 million/m. The strong impinging shock is generated by a wedge of angle 30.960 to the freestream. The shock is made to impinge on a flat plate (made of Hylem which is adiabatic, except for one case with plate made of aluminium which allows heat transfer). The position of (inviscid) shock impingement may be varied (from 55 mm from the leading edge to 100 mm from the leading edge) by moving the plate back and forth on the fixture which holds the wedge and the plate. Expectedly the strong shock generates a large separation bubble of length comparable to the distance of the location of shock impingement from the leading edge of the plate. Such large separation bubbles are typical of supersonic/hypersonic intakes at off-design operation. The evolution of the flow field- including the evolution of impinging shock and subsequent evolution of the large separation bubble- within the short test duration of the shock tunnels is one of the main concerns addressed in the study. Time resolved schlieren flow visualizations using high speed camera, surface pressure measurements using PCB, kulite and MEMS sensors, surface convective heat transfer measurements using platinum thin film sensors are the flow diagnostics used. From the time resolved visualizations and surface pressure measurements with the fast response sensors, the flow field, even with a separation bubble as large as 75 mm (at Mach 5.96, with shock impingement at 95 mm from the leading edge) was found to be established within the short shock tunnel test time. The effects of various parameters- freestream Mach number, distance of the location of shock impingement, freestream total enthalpy and wall heat transfer- on the interaction are investigated. With increase in Mach number from 5.96 to 8.67, for nearly the same shock impingement locations (95 mm and 100 mm from the leading edge respectively), the separation length decreased from 75 mm to 60 mm despite the fact that the shocks are doubly stronger at the higher Mach number. Inflectional trend in separation length was observed with enthalpy at nominal Mach number 8- separation length increased from 60 mm at 1.6 MJ/kg to 70 mm at 2.4 MJ/kg, and decreased drastically to ~40 mm at 6 MJ/kg (when dissociations are expected). The separation length Lsep for all the experiments, except the experiments at 6 MJ/kg, were found to be large, i.e. comparable with the distance xi of location of shock impingement from the leading edge of the flat plate. The scaled separation length (with Hylem wall) was found to obey the inviscid similarity law proposed from the present study for large separation bubbles with strong impinging shocks, where M∞ is the freestream Mach number, p∞ is the freestream pressure and pr is the measured reattachment pressure; this holds for freestream total enthalpy ranging from 1.3 MJ/kg to 2.4 MJ/kg and Reynolds number (based on location of shock impingement) ranging from 1x105 to 4x105. While the increase in separation length from 1.6 MJ/kg to 2.4 MJ/kg could thus be attributed to the small difference in Mach number between the cases (due to inverse variation with cube of Mach number), the decrease in separation length and the non-confirmation to the proposed similarity law for the 6 MJ/kg case is attributed to the real gas effects. At Mach 6 the flow was observed to separate close to the leading edge, even when the (inviscid) shock impingement was at 95 mm from the leading edge. This prompted the proposal of an approximate inviscid model of the interaction for the Mach 6 case with separation at leading edge, and reattachment at the location of (inviscid) shock impingement; Accordingly, the closer the location of impingement, the more the angle that the separated shear layer makes with the plate and hence more the pressure inside the separation bubble. A small reduction in separation length was also observed with aluminium wall when compared with Hylem wall, emphasizing the importance of wall heat conductivity (especially when concerning separated flows) even within the short test durations of shock tunnels. The free interaction theory over adiabatic wall was found to predict the pressure at the location of separation, but under-predict the plateau pressure (at nominal Mach number 8). Numerical simulations (steady, planar) were also carried out using commercial CFD solver FLUENT to complement the experiments. Simulations using one equation turbulence model (Spalart-Allmaras model) were closer to the experimental results than the laminar simulations, suggesting that the flow field may be transitional or turbulent after separation. Significant reduction of the separation bubble length was demonstrated with the control of the interaction using boundary layer bleed within the short test time of the shock tunnel; with tangential blowing at the separation location20% reduction in separation length was observed, while with suction at separation location the reduction was 13.33 %.

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Tseng, Pin-chian, and 曾品蒨. "Study of Shock Waves/Boundary Layer Interactions on Hypersonic Intake Flows." Thesis, 2012. http://ndltd.ncl.edu.tw/handle/04293377258246163738.

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碩士
逢甲大學
航太與系統工程所
100
This paper describes using finite volume method to solve Reynold average Navier-Stokes Equations, simulated a hypersonic intake has a double-ramp compress section. Considering varied geometry to carry out 2-D and 3-D computing simulation of flow field. Computing simulations of the shock wave/boundary layer interaction at the double-ramp and isolator, we have further research about the flow field change result from the 3-D effect. During the research, we test the mesh first, increase the mesh number at the leading edge of the intake, corner between two faces and near wall. We try to use boundary layer mesh, from the result we fine out the boundary layer mesh can effectively reduce mesh number of model without affecting the accuracy of the shock wave simulation. Therefore, we use boundary layer mesh in our computing simulation. About unsteady test, investigate the affect of calculation results with different time step size and time steps, the calculation results show that the time step size is accounted decided of the result, smaller time step size can obtain correct results, but we may cause lengthy computation time if time step size too small. This study further explores the airvent geometry affect of hypersonic intake coupled flow field, the calculated results show the separation point of two ramps corner will delay induced if we increase airvent, and discharge separation bubble induced by the boundary layer smoothly at isolator inlet, improve the flow field in the isolator with shock wave/boundary layer interaction is weakened. The result of 3-D computing simulation show the high-speed airflow will be loss from sides with not consider the sidewall, it can reduce the mass flow entering the engine, 3-D flow field will be improve the phenomenon of loss if we increase sidewall. The corner vortex caused by the sidewall can generate further interaction with separation shock wave, additional shock wave will be induced in coupled flow field of separation bubble.

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48

(6196277), Elizabeth Benitez. "Instability Measurements on Two Cone-Cylinder-Flares at Mach 6." Thesis, 2021.

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This research focuses on measurements of a convective shear-layer instability seen naturally in quiet hypersonic flow. Experiments were carried out in the Boeing/AFOSR Mach 6 Quiet Tunnel (BAM6QT) at Purdue University. The BAM6QT provides low-disturbance hypersonic flow with freestream noise levels similar to what would be experienced by a flight vehicle. To obtain high-speed, off-the-surface measurements of the instability, a modified focused laser differential interferometer (FLDI) was first designed to work with the contoured Plexiglas windows available in the tunnel.

A cone-cylinder-flare geometry was then selected to study the instabilities related to an axisymmetric separation bubble at Mach 6. The sharp cone had a 5-degree half-angle, while flare angles of 10 degrees and 3.5 degrees were tested to compare axisymmetric compression with and without separation, respectively. Under quiet flow, laminar separation and reattachment was confirmed by schlieren and surface pressure-fluctuation measurements. Coherent traveling waves were observed. These were attributed to both the second-mode instability, as well as a shear-generated instability from the separation bubble. The symmetry of the bubble was found to be highly sensitive to angle of attack. Additionally, by introducing controlled disturbances on the cone upstream of the separation, larger-amplitude shear-generated waves were measured while the second-mode amplitudes remained unchanged. Therefore, the shear-generated waves were amplified moving through the shear layer, while the second mode remained neutrally stable. These appear to be the first measurements of traveling waves that are generated in the shear layer of a separation bubble in hypersonic flow.

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49

Lakshman, Srinath. "Experimental Investigations of Leading Edge Bluntness in Shock Boundary Layer Interactions at Hypersonic Speeds." Thesis, 2015. http://etd.iisc.ernet.in/2005/3865.

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Shock Boundary Layer Interactions (SBLIs) and shock-shock interactions are some of the most fundamental problems in high speed aerodynamics. These interactions are of particular importance in scramjet intakes at hypersonic speeds. In hypersonic own with strong SBLI accompanied by own separation, large separation bubbles can form due to high impinging shock strengths. While experiments involving large separation lengths for the impinging shock boundary layer interactions near sharp leading edge are well documented in the literature, only few investigations on the effect of leading edge bluntness on the interactions are studied. In the present study, experiments were carried out to study the role of leading edge bluntness on the impinging shock boundary layer interactions. An oblique shock generated by a wedge (wedge angle 31 degrees) is made to impinge on a at plate (length 200 mm) over which a boundary layer develops. Different leading edge inserts were used on a at plate to get either a sharp or a blunt (radii from 2 to 8 mm) leading edge. The position of the at plate was moved horizontally with respect to the wedge to vary the shock impingement location relative to the leading edge. Experiments were carried out at two freestream conditions - Mach 5.88 (total enthalpy of 1.26 MJ/kg and freestream Reynolds number of 3.85 million per meter) and Mach 8.54 (total enthalpy of 1.85 MJ/kg and freestream Reynolds number of 1.41 million per meter). The various features of the interaction along with different parameters were obtained from schlieren visualizations and surface pressure measurements. The schlieren visualization was used to obtain the separation length, while the reattachment pressure was obtained from the surface pressure distribution. From the present experimental study, a reduction in separation length was observed with an increase in leading edge bluntness. It was also seen that the sharp leading edge had the maximum separation length. Correlations for the separation length and the reattachment pressure have been proposed for these experimental conditions. Numerical simulations were also carried out using commercial software and they had a qualitative agreement with the experiments.

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50

(11022453), Akshay Deshpande. "Unsteady Dynamics of Shock-Wave Boundary-Layer Interactions." Thesis, 2021.

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Shock-wave/turbulent boundary-layer interactions (SWTBLIs) are characterized by low-frequency unsteadiness, amplified aerothermal loads, and a complex three-dimensional flowfield. Presence of a broad range of length and time-scales associated with compressible turbulence generates additional gasdynamic features that interact with different parts of the flowfield via feedback mechanisms. Determining the physics of such flows is of practical importance as they occur frequently in different components of a supersonic/hypersonic aircraft such as inlets operating in both on- and off-design conditions, exhaust nozzles, and control surfaces. SWTBLIs can cause massive flow separation which may trigger unstart by choking the flow in an inlet. On control surfaces, fatigue loading caused by low-frequency shock unsteadiness, coupled with high skin-friction and heat transfer at the surface, can result in failure of the structure.

The objective of this study is twofold. The first aspect involves examining the causes of unsteadiness in SWTBLIs associated with two geometries – a backward facing step flow reattaching on to a ramp, and a highly confined duct flow. Signal processing and statistical techniques are performed on the results obtained from Delayed Detached-Eddy Simulations (DDES) and Implicit Large-Eddy Simulations (ILES). Dynamic Mode Decomposition (DMD) is used as a complement to this analysis, by obtaining a low-dimensional approximation of the flowfield and associating a discrete frequency value to individual modes.

In case of the backward facing step, Fourier analysis of wall-pressure data brought out several energy dominant frequency bands such as separation bubble breathing, oscillations of the reattachment shock, shear-layer flapping, and shedding of vortices from the recirculation zone. The spectra of reattachment shock motion suggested a broadband nature of the oscillations, wherein separation bubble breathing affected the low-frequency motion and shear-layer flapping, and vortex shedding correlated well at higher frequencies. A similar exercise was carried out on the highly confined duct flow which featured separation on the floor and sidewalls. In addition to the low-frequency shock motions, the entire interaction exhibited a cohesive back-and-forth in the streamwise direction as well as a left-right motion along the span. Mode reconstruction using DMD was used in this case to recover complex secondary flows induced by the presence of sidewalls.

For the final aspect of this study, a flow-control actuator was computationally modeled as a sinusoidally varying body-force function. Effects of high-frequency forcing at F⁺ =1.6 on the flowfield corresponding to a backward facing step flow reattaching on to a ramp were examined. Conditionally averaged profile of streamwise velocity fluctuations, based on reattachment shock position, was used for the formulation of spatial distribution of the actuator. The forcing did not change the mean and RMS profiles significantly, but affected the unsteadiness of the interaction significantly. The effects of forcing were localized to the recirculation zone and did not affect the evolution of the shear-layer. The acoustic disturbances propagating through the freestream and recirculation zone drove the motion of the reattachment shock, and did not alter the low-frequency dynamics of the interaction.

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