Experimental Investigation Of The Three Dimensional Flow And Losses In An Axial Flow Turbine Stage

The objective of this investigation is experimental and computational study of three dimensional viscous flow field in the nozzle passage of an axial flow turbine stage. The nozzle passage flow field has been measured using a two sensor hot-wire probe at various axial and radial stations. In addition, two component LDV measurements at one axial station (x/c(sum m) = 0.56) were performed to measure the velocity field. Static pressure measurements and flow visualization, using a fluorescent oil technique, were also performed to obtain the location of transition and the endwall limiting streamlines. A three dimensional boundary layer code, with a simple intermittency transition model, was used to predict the viscous layers along the blade and endwall surfaces. The boundary layers on the blade surface were found to be very thin and mostly laminar, except on the suction surface downstream of 70% axial chord. Strong radial pressure gradient, especially close to the suction surface, induces strong cross flow components in the trailing edge regions of the blade. On the end-walls the boundary layers were much thicker, especially near the suction corner of the casing surface, caused by secondary flow. The secondary flow region near the suction-casing surface corner indicates the presence of the passage vortex detached from the blade surface. The corner vortex is found to be very weak. The presence of a closely spaced rotor downstream (20% of the nozzle vane chord) introduces unsteadiness in the blade passage. The measured instantaneous velocity signal was filtered using FFT square window to remove the periodic unsteadiness introduced by the downstream rotor and fans. The filtering decreased the free stream turbulence level from 2.1% to 0.9% but had no influence on the computed turbulence length scale. The computation of the three dimensional boundary layers is found to be accurate on the nozzle passage blade surfaces, away from the end-walls and the secondary flow region. On ...

An investigation of the flow pattern immediately downstream of an annular cascade of nozzle blades designed for vortex-type velocity distribution and hub discharge angle of approximately 65 degrees from axial was conducted in order to obtain a better understanding of the processes by which the distribution of losses and deviations from design turning angle occur. For the two Mach numbers investigated, the loass region at the outer shroud decreased while that at the inner shroud increased with increasing Mach number.

This paper presents an exhaustive experimental documentation of the three-dimensional nature of the flow in a one-and-one-half stage axial turbine. The intent was to examine the flow within, and downstream of, both the stator and rotor airfoil rows so as to delineate the dominant physical mechanisms. Part 1 of this paper presented the aerodynamic results. Part 2 presents documentation of the mixing, or attenuation, of a simulated spanwise inlet temperature profile as it passed through the turbine, including: (1) the simulated combustor exit-turbine inlet temperature profile, (2) surface measurements on the airfoils and endwalls of the three airfoil rows, and (3) radial-circumferential distributions downstream of each airfoil. Although all three rows contributed to profile attenuation, the impact of the rotor was strongest.

Measurements were made in a large scale single stage turbine facility. Within the nozzle passage measurements were made using a five hole probe, a two-component Laser Doppler Velocimeter (LDV), and a single sensor hot wire probe. These measurements showed weak secondary flows at midchord, and two secondary flow loss cores at the nozzle exit. The casing vortex loss core was the larger of the two. At the exit radial inward flow was found over the entire passage, and was more pronounced in the wake. Nozzle wake decay was found to be more rapid than for an isolated vane row due to the rotor's presence. The midspan rotor flow field was measured using a two-component LDV. Measurements were made from upstream of the rotor to a chord behind the rotor. The distortion of the nozzle wake as it passed through the rotor blade row was determined. The unsteadiness in the rotor flow field was determined. The decay of the rotor wake was also characterized. Zaccaria, M. and Lakshminarayana, B. Glenn Research Center NASA-CR-4778, NAS 1.26:4778, E-10771 NAG3-555; RTOP 523-26-33...