Flow Behavior in a Transonic Axial Compressor Stage

Abstract

The steady and unsteady flow characteristics typically vary along and across the axial compressor stage. This coupled with asymmetric rotor tip clearance that occurs in practice makes flow even more complex. Understanding the complex flow behavior inside the transonic compressor stage will aid in developing flow control devices that are meant for purposes such as improving the rotating stall margin, flutter margin, etc. Here, a detailed time averaged numerical analysis is performed on the single stage transonic axial compressor with averaged rotor tip clearance (1.75% of rotor tip axial chord). An attempt is made to study the compressor stall phenomenon. Computational Fluid Dynamics (CFD) helps in resolving the complex flow features involved in a turbomachinery component and at transonic Mach numbers fairly well. Commercial tool ANSYS CFX is used for solving the 3D compressible Reynolds Averaged Navier-Stokes (RANS) equation with Shear Stress Transport (SST) turbulence model. Grid independency is carried out for three different mesh size models. All mesh models chosen have fine mesh near wall boundary regions to capture the boundary layer effects. Overall performance maps of the compressor are generated for 50% to 100% rated design speeds in steps of 10% for the chosen optimum grid. Flow variations along the blade annulus are studied for three different operating conditions: choke/free flow, peak efficiency and near stall flow conditions and for different speeds. Flow parameters such as Mach number, static and total pressure variations, etc. are studied at the inlet to rotor, exit to rotor and exit to stator for the various flow conditions and speeds. The boundary layer growth is clearly captured when the flow is throttled from choke/free flow conditions to near stall condition for all the speeds investigated. Mach number variation along blade height clearly shows decrease in Mach number as stall is approached. Blade loading distribution of the rotor at hub, mean and tip sections are clearly captured. Shock motion from around mid-chord region at free flow condition to towards the leading edge at near stall condition is clearly highlighted. Velocity streamlines near the tip section show the complex interaction of the tip leakage and clearance flows. Velocity vectors near the blade tip shows, the backflow near the trailing edge and tendency for leading edge spillage as the back pressure is increased. The flow blockage region is captured in the meridional plot and the motion of vortex core region as stall is approached is demarcated in the r-θ plots. Tangential velocity variation across the annulus for the two flow conditions investigated shows stall initiating from the tip section of the blade as compressor is throttled. Flow compensation at near stall conditions is explained.