Time-Accurate CFD Analysis on the Effects of Upstream Stator Loading and Blade Row Interactions on Stator Suction Side Boundary Layer Behavior

Multiple high-fidelity time-accurate computational fluid dynamics simulations were performed to investigate the origin of a low-velocity region at the stator trailing edge in a single stage transonic compressor. Two loadings on the upstream stator row of decreased and nominal were studied at mid spacing of the Blade-Row Interaction Rig. PTURBO was used to generate periodic, quarter annulus simulations of the swirler, deswirler, and rotor blade rows of the BRI compressor rig. Previous results showed that vortex size and strength increase with stator loading due to a large low-velocity region that formed on the stator suction side near the trailing edge. This caused stronger vortices to be shed from the stator trailing edge due to the passing rotor bow shock. At nominal loading the suction side boundary layer thickness increased near 10-20% stator chord due to reflecting pressure waves through the stator passage, creating a low-velocity bubble far upstream on the stator. This low-velocity bubble grew larger as it passed through reflected pressure waves until reaching the trailing edge. The pressure side pressure wave was also observed to have an effect on the pressure gradient at the stator leading edge, increasing the suction side boundary layer thickness just aft of the leading edge. At decreased loading the suction side boundary layer thickness increased much farther downstream, near 55% stator chord, leading to a much smaller low-velocity region, and therefore weaker shock-induced vortices. An understanding of the unsteady interactions associated with blade loading and rotor shock strength in transonic stages will help compressor designers account for unsteady flow physics at design and off-design operating conditions. IGH performance turbomachines typically are designed with highly loaded blade rows with decreased axial spacing, thereby exhibiting significant unsteady losses between blade rows not observed in low-speed turbomachines. These blade-row interactions, such as the interaction of a shock with a blade surface or a blade wake, are a significant source of unsteadiness in high-speed turbomachines. Most contemporary compressor design tools do not directly account for these significant unsteady effects. A better understanding of such phenomena is needed to identify the impact of unsteady aerodynamics on compressor performance, to develop and validate tools for measuring and modeling unsteady flows and to develop design tools that more accurately account for unsteady aerodynamics. Three dimensional experiments and computational simulations are necessary to accurately predict compressor performance, especially in the transonic regime where the rotor leading edge shock accounts for the majority of the pressure rise and loss. Adamczyk 1 described the need for experimental and numerical work which focused on unsteady fluid mechanics and the impact on axial turbomachinery performance. Experimental results increase understanding of these unsteady flows and can also be used to verify results obtained from design tools. Adamczyk described a need for multi-stage design tools that do not rely on empirical formulations or data as inputs. He showed that in order to develop design tools that account for unsteady characteristics, a more complete understanding of unsteady flows that are classified as nondeterministic but are not turbulent in nature must be obtained. An example of these unsteady flow characteristics is the shedding of vortices from a blade’s trailing edge. Vortex shedding in turbomachines has been the focus of research for some time. Hathaway et al. 2 observed vortex shedding in fan rotors, which were shown to lead to spanwise redistribution of entropy by Kotidis and Epstein. 3 It has also been observed that the stretching of vortices leads to flow instabilities which resulted in rapid