One of the important purposes of aero-engine high speed compressor design is to decrease its weight. In order to achieve this purpose, it is required to increase the capability of pressure producing in each individual stage of the compressor. The most common way is to use of high pressure aspect ratio blades. These long and thin blades are exposed to serious vibrations in the high speed flow because of the aeroelastic instability. Mechanical designers link adjacent blades by using Mid-span shroud (damper) to decrease the blade destructive vibrations. This dampers cause flow blockage and turbomachine performance loss. In the previous studies, less attention has been done on the effect of damper on blade shocks, trailing edge vortices and near stall condition. In this paper, aerodynamic performance of a compressor with and without mid-span damper has been investigated and compared. On the other hand, the damper effect on the formation and behavior of shock induced separation has been investigated in each two cases. As a result, presence of damper causes an isentropic efficiency reduction. This damper causes 33% pressure loss on the blades in the region of the extent of 2.7% of the blade span around damper. Turbulence due to the presence of damper leads to the distortion of the vortices pattern on training edge.

Keywords: "Transonic compressor", "Mid-span damper", "Shock induced separation", "Vortex pattern", "stall"

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