Low Reynolds Number Effect on Highly Loaded Compressor Stator Cascade

Authors

Arun Joseph  - Institute of Fluid Flow Machinery Polish Academy of Science IMP PAN, Poland
Pawel Flaszynski - Institute of Fluid Flow Machinery Polish Academy of Science IMP PAN, Poland
Piotr Doerffer - Institute of Fluid Flow Machinery Polish Academy of Science IMP PAN, Poland
Michal Piotrowicz - Institute of Fluid Flow Machinery Polish Academy of Science IMP PAN, Poland

Abstract

Fan and compressor designers are constantly being pushed to higher pressure ratios and lower weights. This leads to a considerable increase in the flow velocities relative to the blades with supersonic speed and shock waves in front of and within the blade passages. At higher altitudes, aircraft engine encounters low Reynolds number (Re) resulting in flow separation. Boundary layer on compressor blade surfaces mostly would be laminar state upstream the shock wave, which is more liable to separate than turbulent boundary layer When shock wave interacts with laminar boundary layer it tends to separate resulting rotating stall in blade passage, which affects the performance. Thus, flying at a higher altitude with low Re has a significant influence on engine performance and aerodynamic stability. Extensive research on how to reduce the flow separation in transonic compressors and improve the flow efficiency is still in progress. The highly loaded compressor stages operated at low Re lead to increased viscous losses. Losses in 2D airfoils operating at low Re have been extensively researched in the past, but very little is known about the 3D aerodynamics of a low Re axial compressor. It is known that the 2D losses increase with decreasing Re. The primary cause for these losses is the transition of flow from laminar to turbulent and the subsequent turbulent separation. The Shock Wave Boundary Layer Interaction (SBLI) will be investigated numerically and experimentally. This paper is an investigation of SBLI on highly loaded compressor stator cascade which are operated under a low Reynolds number of 200000 based on blade chord length and an inflow Mach number of 0.9. Upon these requirements, a steady-state Reynolds Averaged Navier Stokes (RANS) simulation using Explicit Algebraic Reynolds Stress Model (EARSM) turbulence model in Fine/Turbo Numeca was carried out for cascade configuration. A test section has been designed to perform an experimental investigation of SBLI on stator cascades. The test section should be reproducing the same flow structure as in cascade with an identical blade configuration. To retain the periodicity in the test section we extract the streamlines from the cascade simulations to design the upper and lower wall of the test section. To check the flow periodicity in the test section traverse plot upstream and downstream the profiles have been used to compare with cascade simulation. The shock location, secondary flows, and corner separation are influenced by the downstream boundary conditions and sidewalls of the test section. A normal shock is generated at the suction side of the profile. The lambda foot size of the shock and the size of the separation bubble are major factors influencing the shock oscillations. The unsteadiness of the shock wave in the blade passage has been investigated numerically using the Unsteady Reynolds Averaged Navier Stokes (URANS) method. The designed test section will be manufactured and mounted at IMP-PAN transonic wind tunnel facility.