摘要 :
The role of streamline curvature-driven favourable pressure gradients in modifying the turbulence structure of a Mach 4.9, high-Reynolds-number (Re θ = 43000) boundary layer is examined. Three pressure gradient cases (β = (dp/dx...
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The role of streamline curvature-driven favourable pressure gradients in modifying the turbulence structure of a Mach 4.9, high-Reynolds-number (Re θ = 43000) boundary layer is examined. Three pressure gradient cases (β = (dp/dx)(δ/τ_w)= 0. 07,-0. 3 and-1. 0) are characterized via particle image velocimetry. The expected stabilizing trends in the Reynolds stresses are observed, with a sign reversal in the Reynolds shear stress in the outer part of the boundary layer for the strongest favourable pressure gradient considered. The increased transverse normal strain rate and reduced principal strain rate are the primary factors. Reynolds stress quadrant events are redistributed, such that the relative differences between the quadrant magnitudes decreases. Very little preferential quadrant mode selection is observed for the strongest pressure gradient considered. Two-point correlations suggest that the turbulent structures are reoriented to lean farther away from the wall, accompanied by a slight reduction in their characteristic size, consistent with previous flow visualization studies. This reorientation is more pronounced in the outer, dilatation-dominated region of the boundary layer, whereas the alteration in structure size is more pronounced nearer the wall, where the principal strain rates are larger. In addition, integration of a simplified form of a Reynolds stress transport closure model provided a framework to assess the role of the strain-rate field on the observed Reynolds shear stresses. Given the simple geometry, the present data provide a suitable test bed for Reynolds stress transport and large-eddy model development and validation.
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Numerical and analytical investigation of favourable pressure gradient influence on the characteristics of dynamic boundary layer is presented in this work. It is shown that at certain value of pressure gradient in the laminar bou...
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Numerical and analytical investigation of favourable pressure gradient influence on the characteristics of dynamic boundary layer is presented in this work. It is shown that at certain value of pressure gradient in the laminar boundary layer asymptotic conditions are reached, when skin-friction coefficient has no dependence on Reynolds number as at asymptotic suction of the boundary layer through the porous wall. There is some analogy between these two cases of the flow: both at the suction through the porous wall and in the boundary layer of the accelerated flow there is cross-flow directed from the outer boundary to the wall.
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Aerodynamic mixing of a Mach 2 jet controlled with rectangular flat tab with length equal to the nozzle exit diameter, placed at locations 0.25D, 0.5D and 0.75D, downstream of the nozzle exit, has been studied in the presence of d...
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Aerodynamic mixing of a Mach 2 jet controlled with rectangular flat tab with length equal to the nozzle exit diameter, placed at locations 0.25D, 0.5D and 0.75D, downstream of the nozzle exit, has been studied in the presence of different levels of pressure gradients corresponding to nozzle pressure ratio (NPR) range from 3 to 8. The mixing modification associated with shifted tabs is compared with the mixing caused by the same tab at the nozzle exit (OD). The aerodynamic mixing caused by the mass transporting small-scale vortices shed from the edges of the tab placed at the shifted position is found to be appreciably larger than the tab at nozzle exit, for some levels of pressure gradient. For some other levels of nozzle expansion, mixing caused by the shifted tab is comparable to that of tab at nozzle exit. The waves present in the core of the jet controlled with shifted tab were found to be weaker than that of the jet controlled with tab at nozzle exit. At a marginally underexpanded state corresponding to NPR 8, jet core length reduction caused by the tab at 0.75D is about 39.21%, which is closer to the reduction of 40.2%, caused by the tab at OD. The corresponding core length reduction for tab at 0.25D and 0.5D are 38.16% and 20%, respectively.
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A numerical simulation of a square jet ejected transversely into a laminar boundary-layer flow was performed at a jet-to-main-flow velocity ratio of 9.78 and jet Reynolds number of 6330. The jet consisted of a single pulse with a ...
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A numerical simulation of a square jet ejected transversely into a laminar boundary-layer flow was performed at a jet-to-main-flow velocity ratio of 9.78 and jet Reynolds number of 6330. The jet consisted of a single pulse with a duration equal to the time required for the jet fluid to travel 173 jet widths. A strongly-favourable streamwise pressure gradient was applied to the boundary layer and produced a freestream acceleration that is above the typical threshold required for relaminarization. The results of the simulation illustrate the effect of the favourable streamwise pressure gradient on the flowfield created by the transverse jet. Notably, the horseshoe vortex system created upwind of the jet remains steady in time and does not induce noticeable fluctuations in the jet flow. The upwind and downwind shear layers of the jet roll-up through a Kelvin–Helmholtz-like instability into discrete shear-layer vortices. Jet vorticity in the upwind and downwind shear layers accumulates near the corners of the jet and produces two sets of vortex pairs, the former of which couple with the shear-layer vortices to produce large, counter-rotating vortices in the freestream, while the latter are unstable and periodically produce hairpin vortices in the main-flow boundary layer and elongated vortices in the freestream behind the jet. The departure of the jet flowfield from the vortical structures typically observed in transverse jets illustrates the substantive effect of the favourable streamwise pressure gradient on the flowfield created by the jet.
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Shock-wave/boundary-layer interactions (SWBLI) strongly influence the aerodynamic behaviour of many aerospace, propulsion and transportation systems. They have therefore been studied intensively - with different methods, configura...
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Shock-wave/boundary-layer interactions (SWBLI) strongly influence the aerodynamic behaviour of many aerospace, propulsion and transportation systems. They have therefore been studied intensively - with different methods, configurations, and test facilities. Regions of favourable pressure gradient are often placed downstream of the SWBLI region., e.g. to prevent wind-tunnel blockage. To be able to better understand their effect and compare cases from varying sources, a joint experimental-numerical study of a supersonic com-pression/expansion corner flow is carried out. A high-resolution long-time numerical simulation complements particle-image velocimetry measurements, both using the same configuration and conditions. Derived setups without the downstream expansion corner are also explored; both for the same compression-corner angle and for a ramp angle inducing a separation bubble of the same size as for the original compression/expansion configuration. Results demonstrate that the topology and very low-frequency dynamics of the separation bubble developing over the compression corner are altered by the downstream expansion corner. Differences are mainly found for the downstream motion of the shock/shrinking of the separated region and extend downstream of the reattachment point up to the expansion corner.
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