Wake affected boundary layer flows within an embedded stage of a multi-stage axial compressor

Abstract

Boundary Layer Flows in turbomachines have been recognised to crucially influence the stability and performance of gas turbine components particularly the compressor since its function is to provide a pressure rise through diffusion that is accompanied by an adverse pressure gradient. This brings about the danger of separation of flow and thus understanding the behaviour and control of the boundary layer is of great importance. The interaction between stationary and rotating blade rows inevitably make the flow environment within a multi-stage axial compressor unsteady. Dean (1959). Stator blades are subjected to periodic wakes from upstream rotor rows, which initiate transition that is very much unlike those encountered in a steady state (clean) flow conditions. This thesis describes the third stage of the fundamental research on boundary layer flows in turbomachinery at Cranfield University. The feature that makes this experimental work unique is that with constant temperature hot-wire anemometry (CTA) as the principal tool, detailed boundary layer surveys have been carried out within real turbomachinery environment of Cranfield University’s world renowned four-stage low-speed research compressor facility (LSRC). This is a considerable step from the flat plate and wind tunnel arrangement employed in the previous stages of this research programme. Two-dimensional flow field measurements taken downstream of an embedded rotor stage complement the boundary layer survey experiment to contribute toward a better understanding of the unsteady wake affected boundary layer flows that occur in turbomachinery. The current study was conducted with particular reference to a low speed 3 D end bend controlled diffusion blade design. X-array hot-wire measurements of the unsteady flow field have captured the strong wake features exhibiting periodic variations in thickness, which indicates blade loss fluctuations. This is more prominent at Near Stall conditions. Strong variations are present in the end-wall regions where comer stall and comer vortex and tip clearance flows dominate. Radial migration of flow deviation was also detected with each passage of a rotor wake indicating significant radial asymmetries in velocity profiles. Periodic fluctuation of the suction surface boundary layer on a stator blade has been detected starting from the first leading edge measurement location (i.e. 15%) until 37% chord where imminent separation was evident at design flow conditions. Different states of boundary layer flows exist at different heights of a fixed chord position, indicating that the inlet conditions to the stator vary throughout the span. This periodic alternating between laminar and turbulent of the boundary layer appeared to be primarily dependent on the state of the boundary layer itself rather than the wake passing frequency. Results from this first ever attempt of detailed boundary layer survey within turbomachinery environment at Cranfield, though qualitative, has shown that transitional flows occupy up to approximately 40% of the stator blade suction surface. Total separation of the flow, which had been detected at 65% chord in this investigation was not anticipated as the design of the current blading was aimed at delaying if not eliminating any separation until or near the trailing edge.