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

Davis, David O.

doi:10.1115/ajkfluids2015-6342

Cited by 13 publications

(16 citation statements)

References 13 publications

(16 reference statements)

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“…The surface pressure increases rapidly to the peak immediately after a shock wave and then decays in a downstream region of 100 mm. The ratio between the PSP-measured surface pressures after and before a shock wave is p 2 /p 1 ≈ 1.6 − 1.75 in a range of Re D from 1.5 to 5 million, which is consistent with the value of p 2 /p 1 ≈ 1.7 given by Davis [51] based on the pressure tap measurement.…”

Section: Incident Swblisupporting

confidence: 88%

“…A shock generated by a 13.5° cylindrical cone impinges on the wall and interacts with the floor boundary layer at different total pressures and Reynolds numbers ( Re D ) based on the test section diameter. Figure 10 shows the SWBLI region of interest in the Mach 2.5 Axisymmetric test section [50,51], where the origin of the x-coordinate is set at the location of the maximum surface pressure gradient in the surface pressure rise induced by the impinging shock wave. Unsteady PSP measurements were made by using Innovative Scientific Solutions Inc. (ISSI) Turbo PSP with 2 kHz time response.…”

Section: Incident Swblimentioning

confidence: 99%

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Reconstruction of skin friction topology in complex separated flows

Liu

2023

Adv. Aerodyn.

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This paper describes a theoretical method for reconstruction of the skin friction topology in complex separated flows, which is developed based on the exact relation between skin friction and surface pressure through the boundary enstrophy flux (BEF). The key of this method is that a skin friction field is reconstructed from a surface pressure field as an inverse problem by applying a variational method. For applications, the approximate method is proposed, where the composite surface pressure field is given by a linear superposition of the base-flow surface pressure field and the surface pressure variation field and the base-flow BEF field is used as the first-order approximation. This approximate method is constructive in a mathematical sense since a complex skin friction field in separated flows can be reconstructed from some elemental skin friction structures (skin friction source/sink, vortex and their combinations) by a linear superposition of some simple surface pressure structures. The distinct topological features, such as critical points, separation lines and attachment lines, naturally occur as a result of such reconstruction. As examples, some elemental skin friction structures in separated flows are reconstructed in simulations, and the skin friction fields in shock-wave/boundary-layer interactions (SWBLIs) are reconstructed from pressure sensitive paint (PSP) images obtained in wind tunnel experiments.

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Section: Incident Swblisupporting

confidence: 88%

Section: Incident Swblimentioning

confidence: 99%

Reconstruction of skin friction topology in complex separated flows

Liu

2023

Adv. Aerodyn.

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“…In a previous study in the GRC 225 cm 2 SWT, 8 the boundary layer and displacement thickness at three axial locations in the test section were measured at a ReD of 4.0 x 10 6 . The measurements are shown below in Table 1.…”

Section: A) B)mentioning

confidence: 99%

Blockage Testing in the NASA Glenn 225 Square Centimeter Supersonic Wind Tunnel

Sevier

Davis

Schoenenberger

2017

55th AIAA Aerospace Sciences Meeting

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A feasibility study is in progress at NASA Glenn Research Center to implement a magnetic suspension and balance system in the 225 cm 2 Supersonic Wind Tunnel for the purpose of testing the dynamic stability of blunt bodies. An important area of investigation in this study was determining the optimum size of the model and the iron spherical core inside of it. In order to minimize the required magnetic field and thus the size of the magnetic suspension system, it was determined that the test model should be as large as possible. Blockage tests were conducted to determine the largest possible model that would allow for tunnel start at Mach 2, 2.5, and 3. Three different forebody model geometries were tested at different Mach numbers, axial locations in the tunnel, and in both a square and axisymmetric test section. Experimental results showed that different model geometries produced more varied results at higher Mach Numbers. It was also shown that testing closer to the nozzle allowed larger models to start compared with testing near the end of the test section. Finally, allowable model blockage was larger in the axisymmetric test section compared with the square test section at the same Mach number. This testing answered key questions posed by the feasibility study and will be used in the future to dictate model size and performance required from the magnetic suspension system.

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“…Some recent work by Zuo et al [26] considered the interaction of a conical shock wave with a flat plate boundary layer. The interaction of a conical shock wave with an axisymmetric boundary layer to produce a purely 2D SBLI has been addressed in only a handful of publications [27,28]. Such interactions are not only highly relevant for circular inlet flows but also help avoid issues faced in SBLI experiments in rectangular test sections [29,30], such as corner effects, which are exacerbated in the presence of crossflow and 3D flow features.…”

Section: Introductionmentioning

confidence: 99%

Numerical Investigation of Asymmetric Mach 2.5 Turbulent Shock Wave Boundary Layer Interaction

Groß

Slater

2023

Aerospace

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

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

Cited by 13 publications

References 13 publications

Reconstruction of skin friction topology in complex separated flows

Reconstruction of skin friction topology in complex separated flows

Blockage Testing in the NASA Glenn 225 Square Centimeter Supersonic Wind Tunnel

Numerical Investigation of Asymmetric Mach 2.5 Turbulent Shock Wave Boundary Layer Interaction

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