Abstract
Streamwise and spanwise flow pattern over three rounded contour bumps with different flow control strategies employed have been experimentally investigated in a Mach 1.3 freestream. Surface oil flow visualisation, Schlieren photography and particle image velocimetry measurements were used for flow diagnostics. Experimental data showed that in a Mach 1.3 freestream over the baseline plain bump, significant flow separation appeared at the bump crest that led to the formation of a large wake region downstream. In addition, two large counter-rotating spanwise vortices were formed in the bump valley. It was observed that the use of the passive by-pass blowing jet in the bump valley showed no obvious effects in reducing the sizes of both the wake region and the spanwise vortices in the bump valley. In contrast, it was found that the size of the wake region and the spanwise vortices could be reduced by blowing sonic jet in the bump valley. This approach of flow control found to be the most effective when the total pressure of the blowing jet was 2bar. It is deduced that the active blowing jet hindered the formation of the spanwise vortices in the bump valley as well as deflected the shear layer downwards so that a smaller re-circulating bubble was formed downstream of the bump crest.
Highlights
In recent years, a number of studies have been conducted to investigate the potential of implementing rounded contour bumps into aerofoils or aircraft wings in attempt to achieve wave drag reduction in transonic aircraft [1,2,3,4,5]
Understanding the flow physics of rounded contour bumps is important as Konig et al [2,3], Lo et al [6], Bruce and Colliss [7] and Colliss et al [8] have shown that unfavourable flow features like large spanwise counter-rotating vortices and three-dimensional vortical structures could form in the bump valley due to the appearance of flow separation at the bump crest
Three rounded contour bumps known as the plain bump, the passive jet bump and the active jet bump, were used
Summary
A number of studies have been conducted to investigate the potential of implementing rounded contour bumps into aerofoils or aircraft wings in attempt to achieve wave drag reduction in transonic aircraft [1,2,3,4,5]. Understanding the flow physics of rounded contour bumps is important as Konig et al [2,3], Lo et al [6], Bruce and Colliss [7] and Colliss et al [8] have shown that unfavourable flow features like large spanwise counter-rotating vortices and three-dimensional vortical structures could form in the bump valley due to the appearance of flow separation at the bump crest. The presence of these unfavourable flow features and the occurrence of flow separation in the bump crest increase the pressure drag encountered by the contour bumps. The drag reduction gained by installing contour bumps on to the wings could be par-
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