The influence of small cylindrical bluntness of the leading edge of a flat plate on formation of spatial structures in a nominally two-dimensional supersonic compression corner flow at the Mach number M∞ ≈ 8 and a laminar state of the undisturbed boundary layer is studied by the method of temperature-sensitive paints. Streamwise vortices are found in the region of reattachment of the separated flow in a wide range of Reynolds numbers (0.15 · 106–2.55 · 106) for various angles of flow deflection and plate lengths. It is demonstrated that the existence of these vortices induces spanwise oscillations of the heat transfer coefficient; the amplitude of these oscillations may reach 30%. The maximum deviations of the Stanton number reaching 80% are observed in the case with significant roughness of the leading edge of the flat plate. Both the maximum Stanton numbers in the reattachment region and the amplitude of spanwise oscillations of the Stanton number induced by streamwise vortices are found to decrease significantly in the case of small bluntness of the leading edge. Solutions of three-dimensional Navier–Stokes equations are obtained for some test conditions. The computed results are in good agreement with experimental data, which points to a significant stabilizing effect of small bluntness on the intensity of streamwise vortices.

3d secondary separation at supersonic flow over the compression ramp

Effect of Small Bluntness on Formation of Görtler Vortices in a Supersonic Compression Corner Flow

http://enu.kz/repository/2011/AIAA-2011-730.pdf

Three-dimensional Flow Structures at Supersonic Flow over the Compression Ramp

Incipient separation in laminar, ramp-induced shock-wave/boundary-layer interactions is examined using numerical solutions of the triple-deck equations and the full set of governing equations for c...

Incipient Separation in Laminar Ramp-Induced Shock-Wave/Boundary-Layer Interactions

by Brock Edwin Hedlund This thesis documents the efforts to characterize the Mach 4.5 flow over a flat plate model with a 30° compression ramp over a range of unit Reynolds numbers (Rel = 4‧10 5 – 1‧10 m) in low enthalpy (T0 = 293 K) and high enthalpy (T0 = 800 – 1250 K) flow conditions. The compression ramp model was designed, fabricated, and implemented in a hypersonic test facility (ACT-1) to simulate small-scale scramjet inlet operation. Flow visualization was provided by a schlieren imaging system and planar laser Rayleigh scattering (PLRS) to characterize the flow structure with various Rel. Three measurement methods were used to determine the frequency spectrum of flow perturbations: a highfrequency Shack-Hartmann wavefront sensor, a laser differential interferometer, and highfrequency pressure sensors. All three methods provide complementary results to indicate the existence of dominant frequencies of perturbations in the flow. It is determined that of these three measurement tools, Shack-Hartmann wavefront sensing is the most beneficial for this flow analysis. These measurements detect modification to the flow perturbation

3d secondary separation at supersonic flow over the compression ramp

Citations

Effect of Small Bluntness on Formation of Görtler Vortices in a Supersonic Compression Corner Flow

Three-dimensional Flow Structures at Supersonic Flow over the Compression Ramp

Incipient Separation in Laminar Ramp-Induced Shock-Wave/Boundary-Layer Interactions

Measurement of Flow Perturbations Occurring Both Naturally and by Plasma Induction over a Mach 4.5 Corner Separation Zone

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