Top PDF Shock Wave Interactions: A CFD Study of CUBRC LENS-II Turbulent Experiments

Shock Wave Interactions: A CFD Study of CUBRC LENS-II Turbulent Experiments

Shock Wave Interactions: A CFD Study of CUBRC LENS-II Turbulent Experiments

Sufficient run length to ensure natural transition ahead of flare (interaction region) 7 ° cone is identical to that of HIFiRE-1 configuration. HIFiRE-1 had a cylindrical section before [r]

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Optimal Control of Shock Wave Turbulent Boundary Layer Interactions Using Micro-Array Actuation

Optimal Control of Shock Wave Turbulent Boundary Layer Interactions Using Micro-Array Actuation

This research study on micro-actuator array flow control is a collaborative effort between the AFRL Flight Vehicles Directorate (AFRL/VAAI) and NASA Glenn Research Center (NASA/GRC) to develop and demonstrate advanced flow control techniques for supersonic mixed compression inlets. The AFRL Flight Vehicles Directorate has focused recent effort on the advancement of propulsion integration technologies associated with LRSA (Long Range Strike Aircraft). The goal of this effort has been to define and develop high payoff advanced technology approaches for mixed compression inlets operating in the Mach 2.0 to 4.0 speed regime. In addition, NASA/GRC is also focusing a portion of its micro-flow control research effort on commercial and business type aircraft in the speed range Mach 1.6 to 2.0. Therefore, the collaborative effort also provides technology interchange between the military and commercial side of aerodynamic research. The collaboration between AFRL/VAAI and NASA/GRC combines CFD analysis, Design-of-Experiments (DOE) methodologies (4) and experimental test, from small scale subcomponent applications to low-cost DOE based large scale isolated inlet models, to demonstrate reliable high inlet performance as well as unprecedented efficiency of the design/development process. The matured plan for collaboration will involve CFD capabilities in NASA, AFRL, industry and academia. NASA/GRC has developed the baseline CFD analysis and DOE investigation of advanced flow control “fail safe” actuators along with “Proof-of- Concept” CFD analysis of the integration of micro-actuator devices into “real” supersonic external and mixed compression inlet systems. The AFRL/VA Computational Branch (VAAC) has become involved as well, helping to guide test planning at AFRL by providing CFD analysis of wind tunnel hardware. In a collaborative program with NASA/GRC, they are also assisting in providing comparisons of actuator CFD analysis between the NASA WIND code and the AFRL AVUS code using formal statistical analysis. Supplemental basic research into the physics of SWBL interactions will be accomplished by the Univ. of Illinois (CFD Analysis) and Cambridge Univ. (Experimental Studies). Their combined work will provide insight into the fundamental aerodynamic characteristics of “non-bleed” shock wave boundary layer control, both with regards to performance and inlet stability
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CFD designed experiments for shock wave/boundary layer interactions in hypersonic ducted flows

CFD designed experiments for shock wave/boundary layer interactions in hypersonic ducted flows

For high speed air-breathing engines, knowledge of the point at which boundary layer separation occurs limits the design pa- rameters. Shock wave/turbulent boundary layer interactions are a common occurrence in supersonic flows with almost any flow deflection accompanied by shock formation. Incident shock in- teractions occur when the shock that impinges on the boundary layer is generated by an external source. These allow for the study of the interaction of bulk flow compression without the added effects of streamline curvature and hence they have been used for the experimental work in this paper. They are particu- larly important for scramjet studies which involve ducted flows where there is a requirement to add as much heat and pressure as possible. Unfortunately analytical means of modeling sep- arated flow are not advanced. CFD codes however have pro- gressed significantly to the point where several commercially available codes are capable of simulating hypersonic flows in reasonable time frames. When dealing with separated flows it is important to ensure the use of time accurate codes to capture upstream influences which is not possible with time marching codes. Turbulence models still need to be employed to approxi- mate turbulent effects and these are most probably the cause of a large proportion of inaccuracies. Choice of the most appropri- ate turbulence model is therefore very important. Two-equation models are far more accurate when predicting boundary layer separation[1] however for unseparated flows simple algebraic
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Instabilities in oblique shock wave/laminar boundary-layer interactions

Instabilities in oblique shock wave/laminar boundary-layer interactions

The interaction of an oblique shock wave and a laminar boundary layer developing over a flat plate is investigated by means of numerical simulation and global linear-stability analysis. Under the selected flow conditions (free-stream Mach numbers, Reynolds numbers and shock-wave angles), the incoming boundary layer undergoes separation due to the adverse pressure gradient. For a wide range of flow parameters, the oblique shock wave/boundary-layer interaction (OSWBLI) is seen to be globally stable. We show that the onset of two-dimensional large-scale structures is generated by selective noise amplification that is described for each frequency, in a linear framework, by wave-packet trains composed of several global modes. A detailed analysis of both the eigenspectrum and eigenfunctions gives some insight into the relationship between spatial scales (shape and localization) and frequencies. In particular, OSWBLI exhibits a universal behaviour. The lowest frequencies correspond to structures mainly located near the separated shock that emit radiation in the form of Mach waves and are scaled by the interaction length. The medium frequencies are associated with structures mainly localized in the shear layer and are scaled by the displacement thickness at the impact. The linear process by which OSWBLI selects frequencies is analysed by means of the global resolvent. It shows that unsteadiness are mainly associated with instabilities arising from the shear layer. For the lower frequency range, there is no particular selectivity in a linear framework. Two-dimensional numerical simulations show that the linear behaviour is modified for moderate forcing amplitudes by nonlinear mechanisms leading to a significant amplification of low frequencies. Finally, based on the present results, we draw some hypotheses concerning the onset of unsteadiness observed in shock wave/turbulent boundary-layer interactions.
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A Numerical Study of the Limiting Cases of Cylinder-Induced Shock Wave/Boundary Layer Interactions

A Numerical Study of the Limiting Cases of Cylinder-Induced Shock Wave/Boundary Layer Interactions

Since the flow above the upper surface becomes tangential to the surface, the upper boundary above the flat plate may follow suit. From the flow solution for the domain of Figure 3.3 it was determined that despite the far-field condition being implemented, the flat plate leading edge shock had not dissipated out by the time it reached the upper boundary, and was reflecting back into the domain. This reflection impinged onto the flat plate boundary layer. In order to mitigate this issue, far-field absorbing layers were applied to the upper boundary. These posed the only contribution (other than the turbulence model) to the source term vector, Ṡ, in the RHS of Equation (3-1). Physically, the absorbing layers acted as a sponge boundary, where the flow was damped to user-specified freestream values over the course of several layers. The shock dissipated within these damping layers and did not reflect back into the domain, thus allowing for the implementation of the domain in Figure 3.6. The walls here were created with y + = 0.1 and GR = 1.1; the fineness was required to properly resolve the small boundary layer in the vicinity of the leading
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Design and Characterization of a Supersonic Wind Tunnel for the Study of Shock Wave Boundary Layer Interactions

Design and Characterization of a Supersonic Wind Tunnel for the Study of Shock Wave Boundary Layer Interactions

authority of the LAFPAs is most likely dependent on interaction strength. To date, only the effects of frequency, streamwise location, and mode of operation on the LAFPA’s control authority have been studied. The mode of operation has not been thoroughly investigated, as only two modes have been tested. A frequency sweep of the actuators including Strouhal numbers of 0.03 and 0.5 needs to be performed within the new tunnel; previous results suggest that a Strouhal number of 0.03 (the low-frequency unsteadiness present in the reflected shock of an unforced SWBLI) will yield the greatest control authority. In jet exhaust experiments, the LAFPAs have shown an increased effectiveness as the duty cycle is decreased (as long as complete breakdown occurs). The variable angle wedge allows for controllable interaction strength, so an investigation of the dependence of LAFPAs control authority on interaction strength is necessary.
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Transitional shock-wave/boundary-layer interactions in hypersonic flow

Transitional shock-wave/boundary-layer interactions in hypersonic flow

The transition process initiated with linear instabilities is only one of several routes through to turbulence, for which Reshotko (2001) provided a revised version of Morkovin’s road map including transient growth approaches to bypass transition. DNS of bypass transition at low speed have been shown, for example by Jacobs & Durbin (2001), to include turbulent spots. Similar spots have been observed previously in high-speed flows (Fiala et al. 2006) and their growth rates measured in DNS (Redford, Sandham & Roberts 2012). For transitional flows, particularly those with higher background disturbances, the range of phenomena between fully laminar and fully turbulent interactions may be described using the idea of boundary-layer intermittency, introduced by Narasimha (1985), based on the idea of the transition process being dominated by the growth and merging of turbulent spots. An intermittency factor γ is introduced that can give a measure of the state of the boundary layer at a point of shock interaction, equal to zero in laminar flow and one in turbulent flow. There are multiple definitions of intermittency that can be employed based on what can be easily measured (Hedley & Keffer 1974). Common definitions involve the application of a threshold to a particular quantity, usually sampled over a window in time or space. Where the threshold is exceeded, turbulent flow is assumed, otherwise the flow is assumed to be laminar. An alternative approach was proposed by Schneider (1995) based on plots of the probability density function (p.d.f.) of a relevant quantity, for example the skin friction. For his experiments, Schneider found two peaks in the p.d.f. that could be attributed to the laminar and turbulent regions of flow (outside and inside turbulent spots, respectively), leading to a method of defining intermittency without an explicit threshold. In addition to its use in quantifying the transition process, intermittency has been found to be an effective parameter in the development of transition models (see, for example, Steelant & Dick 1996; Langtry & Menter 2009; Steelant & Dick 2001).
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Shock Wave Turbulent Boundary Layer Interaction in a 2-D Compression Corner

Shock Wave Turbulent Boundary Layer Interaction in a 2-D Compression Corner

A computational study has been carried out to analyze the supersonic shock wave turbulent boundary layer interaction in a 2-D compression corner for a free stream Mach number of 2.94. The study has been done for a unit Reynolds number of 36.4x 10 6 per meter and 20 0 corner angle. The model has been analyzed using 2-D numerical simulations based on a commercially available Computational Fluid Dynamics (CFD) Code that employs k-ω Shear Stress Transport (SST) turbulence model. The substantiation of the CFD code and the turbulence model used is obtained by comparing with the experimental results available in literature. Comparison of the surface pressure distribution with experiment exhibited good engineering agreement. Numerical results indicate that the extent of the separated zone has increased and thus show increased separation and reattachment points when compared to experiment.
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Implicit CFD methods for transitional shock wave: boundary layer interaction

Implicit CFD methods for transitional shock wave: boundary layer interaction

Lee (2001) presented a comprehensive review of shock-buffet studies, including physical models of the shock-buffet mechanism. These mechanisms were discussed for symmetrical circular-arc aerofoils at zero incidences and for supercritical aerofoils at high incidence angles with fully separated flows. During the examinations, Lee suggested that the buffet period is equal to the sum of the downstream and upstream wave’s motion time. Those pressure waves are produced either from the shock wave moving downstream to the trailing edge or the trailing edge boundary layer and move upstream towards the shock. Estimating these motions by empirical evaluation over a single buffet cycle, Lee obtained good agreement with experiments.
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Effects of micro-ramps on a shock wave/turbulent boundary layer interaction

Effects of micro-ramps on a shock wave/turbulent boundary layer interaction

To consolidate the major observations made, a conceptual model of the effects of micro-ramps on the SWTBLI is now presented. In the model, the micro-ramps generate on the mean flow level longitudinal streamwise vortices that induce low-speed regions downstream of vertex locations and high-speed regions at intermediate spanwise locations. Instantaneously, however, no streamwise vortices are apparent. Instead, individual structures are convected within the boundary layer, which consist of pairs of counter-rotating vortices. The observed presence of a counter-rotating vor- tex pair is reminiscent of the conditional eddies obtained by Tomkins and Adrian [30], who identified them as cross- sections through conditional hairpin vortices. The generation of hairpin vortices by objects in our boundary layer would be consistent with results reported for incompressible bound- ary layers. For instance, Tufo et al. [32] have found trains of hairpin vortices in the wake of a hemispherical roughness element in a low-speed turbulent incompressible boundary layer using direct numerical simulations (DNS), similar as in the experimental investigations of Acerlar and Smith [33], and we anticipate that a comparable phenomenology also exists downstream of our micro-ramp SBVGs. The concep- tual model is schematically summarized in Fig. 11a. As a result of this modification within the incoming boundary layer, downstream of a low-speed streak in the interaction region, the flow becomes sonic farther upstream. Conversely, the opposite trend occurs downstream of a high-speed region. Thus, the mean flow organization of the subsonic region con- forms to the mean velocity distribution within the incoming boundary layer. Figure 11b shows a conceptual sketch of the perturbed interaction in the mean flow sense. Note that both drawings are not shown to scale.
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Flow instability in shock tube due to shock wave-boundary layer-contact surface interactions, a numerical study

Flow instability in shock tube due to shock wave-boundary layer-contact surface interactions, a numerical study

This paper describes numerical simulation of transient flow conditions in shock tube. A two dimensional time accurate Navier-Stokes CFD solver for shock tube applications is developed to perform the numerical investigations. The solver was developed based on the dimensions of a newly built short-duration high speed flow test facility at Universiti Tenaga Nasional “UNITEN” in Malaysia. The facility has been designed, built, and commissioned in such a way so that it can be used as a free piston compressor, shock tube, shock tunnel and gun tunnel interchangeably. Different values of diaphragm pressure ratios P 4 /P 1 are applicable in order to get wide range of Mach number.
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Experiments on Shock Induced Laminar-Turbulent Transition on a Flat Plate at Mach 6

Experiments on Shock Induced Laminar-Turbulent Transition on a Flat Plate at Mach 6

The infrared pictures of the PEEK insert (figure 8) show that the free transition (without shock generator) is not a straight line figures 8a and 8c. This has to be taken into account when interpreting the results of sensor data which don’t have the same Y-coordinate. This is true for most of the sensor rows used in this project. The waviness of the transition region increases with a decreasing unit Reynolds number. With the shock generator at the front position at the high Reynolds number condition transition occurs more or less on a straight line (figure 8f). This becomes a little wavy again for the middle Reynolds number. In case of the low Reynolds number, the transition region is streaky and has a bean like shape (figure 8d). Probably the transition process is not completed in this case and the boundary layer relaminerizes. The values in figures 8d and 8f with x < 140 mm are not valid, since the view of the camera is blocked by the shock generator there. In analogy to that values with x < 240 mm in figures 8g and 8i are invalid. In contrast to the runs with free transition and the shock generator at the front position the highest Reynolds number does not lead to the highest but to the lowest Stanton numbers when the shock generator is at the back position. In figure 8i also the impingement of the reflected shock from the leading edge is visible at x ≈ 400 mm.
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CFD Validation Experiment of a Mach 2.5 Axisymmetric Shock-Wave/Boundary-Layer Interaction

CFD Validation Experiment of a Mach 2.5 Axisymmetric Shock-Wave/Boundary-Layer Interaction

In order to avoid the pitfalls of a rectangular configuration, an axisymmetric configuration is proposed that is two- dimensional in the mean. The selected interaction is illustrated in Figure 2. A Mach 2.5 core flow approaches a cone-cylinder centerbody that generates a conical shock that impinges and reflects off the cylindrical test section wall, interacting with the naturally occurring test section boundary layer. The approximate measurement area of interest is indicated by the rectangular box shown in the figure. NASA GRC’s supersonic facilities, however, all have square or rectangular test sections so a new facility has been designed specifically for this study. This configuration is similar to a study performed by Rose (Ref. 6), which was considered for Settles and Dodson’s validation database, but was rejected due to questions about the accuracy of the hot-wire measurements. A new 17 cm diameter axisymmetric supersonic wind tunnel (Axi-SWT) has been installed in Test Cell W6B and replaces the existing 15×15 cm configuration. The facility design allows for relatively easy changes between the square and circular configurations.
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Experimental Investigation of the Three-Dimensional Structure of a Shock Wave/Turbulent Boundary Layer Interaction

Experimental Investigation of the Three-Dimensional Structure of a Shock Wave/Turbulent Boundary Layer Interaction

The streamwise and vertical components of the first eigenmode are shown in figures 6(a, b), respectively, with the modal number and energy content shown inset. One caveat before proceeding: Data for y/δ>0.5 contaminated the eigenmode constructions due to unreliable velocity measurements related to the poor distribution of laser light in this region. Thus, only data for y/δ<0.5 are used in the analysis. Furthermore, as the temporal coefficients associated with each eigenmode for a particular realization can be either positive or negative, only the relative changes of sign within an eigenmode are important. Figure 6(a) portrays relatively large streamwise fluctuations within the incoming boundary layer that are energetically associated with a broad spanwise region of fluctuations within the reflected shock foot. The superposition of these fluctuations onto the mean flow
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Heat transfer amplification in transitional shock-wave/boundary-layer interactions

Heat transfer amplification in transitional shock-wave/boundary-layer interactions

to resolve such small structures surpass the capabilities of modern CFD simulation. Therefore, shock waves have to be simulated on much coarser grids, leading to an extended shock wave thickness in the simulations. In this work, a grid convergence study was conducted with the Menter Shear Stress Transport (SST) turbulence model with a fully turbulent boundary layer and four different grid resolutions were investigated. The computational grid was generated using CENTAUR™. The schematic CFD setup is shown in Fig. 4 for the case of α = 3° and the shock generator mounted at Pos. C. The grid has a relative large ( ≈ 12 mm) prism layer (region A) to ensure the shock induced separation bubble is calculated in the prism layer even for the extreme cases. Because this work focuses on heat fluxes, the wall distance of the first grid cell y + has to be significantly below one 11,12 to get reliable results. In the current
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An investigation of bleed configurations and their effect on shock wave/boundary layer interactions

An investigation of bleed configurations and their effect on shock wave/boundary layer interactions

an oblique incident shock, on alaminar boundarylayer, with nique ofBeam and Warming_ toadvancethesolution, per- bleed through a normal slot across the shock impingement formed best both [r]

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CFD Validation Experiment of a Mach 2.5 Axisymmetric Shock-Wave Boundary-Layer Interaction

CFD Validation Experiment of a Mach 2.5 Axisymmetric Shock-Wave Boundary-Layer Interaction

The impinging duct configuration was chosen for two reasons. First, inasmuch as the intent of the investigation is to provide CFD validation data, this configuration allows for a relatively thick incoming boundary layer so highly resolved measurements are possible. And second, although not intended to mimic any particular application, it is of the same general configuration as a SWBLI occurring on the cowl surface of axisymmetric inlets with supersonic internal compression. Previous investigations of this flow configuration include the development of integral flow models for solid and porous walls by Seebaugh et al. [7]. An experimental investigation by Seebaugh and Childs [8] presented surface static and flowfield Pitot pressure measurements under Mach 2.82 and 3.78 flow conditions with cone angles of 10, 13 and 15°. Rose [6] acquired detailed turbulence measurements using hot-wire anemometry under Mach 3.88 flow conditions with a 9° cone angle. Neither of the latter two studies, however, were considered to meet the criteria for CFD validation purposes as proposed by Ref. [1].
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CFD Validation Experiment of a Mach 2.5 Axisymmetric Shock-Wave/Boundary-Layer Interaction

CFD Validation Experiment of a Mach 2.5 Axisymmetric Shock-Wave/Boundary-Layer Interaction

Background • Experimental investigations of specific flow phenomena, e.g., Shock Wave/Boundary-Layer Interactions (SWBLI), provide great insight to the flow behavior but often lack the necessary details to be useful as CFD validation experiments.

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Detonation driven shock wave interactions with perforated plates

Detonation driven shock wave interactions with perforated plates

A DynoStart 2 Spark generator with a capacitance of 0.2 µF and output voltage of 2500 V was used to initiate the detonation. Both DynoLine and DynoStart were manufactured by Dyno Nobel Sweden AB (now, Orica Mining Services). The DynoLine contains a mixture of Octogen (HMX) (∼92% by weight) and traces of Aluminium (∼8% by weight) at 18 mg/m length of the tube. The energy in these blast waves has been estimated to be about 1.25 J. The shock wave created propagates along the tube and diffracts into the ambient, closely followed by the products of combustion which comprises the reaction front, as shown in Figure 3. Shortly after the shock front leaves the NONEL, except a small region along the tube axis, the shock front decouples from the reaction zone, these zones are labelled in Figures 3(a) and (b). The decoupled reaction zone contains shocked but unreacted gases. 22
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Coupled simulation of shock-wave/turbulent boundary-layer interaction over a flexible panel

Coupled simulation of shock-wave/turbulent boundary-layer interaction over a flexible panel

displacement amplitudes differ significantly from the experiment. The authors believe, that the compressible air within the cavity constitutes an additional stiffness which is not included in our setup, but could be possibly modeled through an additional set of springs acting on the underside of the elastic panel. This would in turn increase the frequency of the panel oscillation and explain the current mismatch between experiment and simulation. Regarding the oscillation amplitudes, a relatively high damping is observed in the experiment, which is not the case for the LES. The authors believe, that two effects are mainly responsible for this discrepancy: The cavity is sealed through a thick layer of a soft foam rubber, which may increase the overall damping of the system and leads to a damping effect that scales linearly with the panel velocity. The second reason is attributed to the aerodynamic damping imposed by the cavity, which is expected to scale with the square of the panel velocity. As it is not yet clear how to model this damping behavior cor- rectly, more experiments are necessary for modeling purposes. Therefore, damping in the structural model is neglected in the current numerical setup.
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