No abstract

Compressible, subsonic flow through a diffusing Sduct has been experimentally investigated. Benchmark aerodynamic data are presented for flow through a representative S-duct configuration. The collected data would be beneficial to aircraft inlet designers and is suitable for the validation of computational codes. Measurements of the three dimensional velocity field and total and static pressures were obtained at five cross-sectional planes. Surface static pressures and flow visualization also helped to reveal flow field characteristics. All reported tests were conducted with an inlet centerline Mach number of 0.6 and a Reynolds number, based on the inlet centerline velocity and duct inlet diameter, of 2.6 x 10 6 . The results show that a large region of streamwise flow separation occurred within the duct. Details about the separated flow region, including mechanisms which drive this complicated flow phenomenon, are discussed. Transverse velocity components indicate that the duct curvature induces strong pressure driven secondary flows, which evolve into a large pair of counter-rotating vortices. These vortices convect the low momentum fluid of the boundary layer towards the center of the duct, degrading both the uniformity and magnitude of the total pressure profile.

This is Part Two of a four-part paper. It begins with Section 6.0 and continues to describe the comprehensive experiments and computational analyses that have led to a detailed picture of boundary layer development on airfoil surfaces in multistage turbomachinery. In this part, we present the experimental evidence used to construct the composite picture for compressors given in the discussion in Section 5.0 of Part 1. We show the data from the surface hot-film gages and the boundary layer surveys, give a thorough interpretation for the baseline operating condition, and then show how this picture changes with variations in Reynolds number, airfoil loading, frequency of occurrence of wakes and wake turbulence intensity. Detailed flow features are described using raw time traces. The use of rods to simulate airfoil wakes is also evaluated.

Comprehensive experiments and computational analyses were conducted to understand boundary layer development on airfoil surfaces in multistage, axial-flow compressors and LP turbines. The tests were run over a broad range of Reynolds numbers and loading levels in large, low-speed research facilities which simulate the relevant aerodynamic features of modern engine components. Measurements of boundary layer characteristics were obtained by using arrays of densely packed, hot-film gauges mounted on airfoil surfaces and by making boundary layer surveys with hot wire probes. Computational predictions were made using both steady flow codes and an unsteady flow code. This is the first time that time-resolved boundary layer measurements and detailed comparisons of measured data with predictions of boundary layer codes have been reported for multistage compressor and turbine blading. Part 1 of this paper summarizes all of our experimental findings by using sketches to show how boundary layers develop on compressor and turbine blading. Parts 2 and 3 present the detailed experimental results for the compressor and turbine, respectively. Part 4 presents computational analyses and discusses comparisons with experimental data. Readers not interested in experimental detail can go directly from Part 1 to Part 4. For both compressor and turbine blading, the experimental results show large extents of laminar and transitional flow on the suction surface of embedded stages, with the boundary layer generally developing along two distinct but coupled paths. One path lies approximately under the wake trajectory while the other lies between wakes. Along both paths the boundary layer clearly goes from laminar to transitional to turbulent. The wake path and the non-wake path are coupled by a calmed region, which, being generated by turbulent spots produced in the wake path, is effective in suppressing flow separation and delaying transition in the non-wake path. The location and strength of the various regions within the paths, such as wake-induced transitional and turbulent strips, vary with Reynolds number, loading level, and turbulence intensity. On the pressure surface, transition takes place near the leading edge for the blading tested. For both surfaces, bypass transition and separated-flow transition were observed. Classical Tollmien–Schlichting transition did not play a significant role. Comparisons of embedded and first-stage results were also made to assess the relevance of applying single-stage and cascade studies to the multistage environment. Although doing well under certain conditions, the codes in general could not adequately predict the onset and extent of transition in regions affected by calming. However, assessments are made to guide designers in using current predictive schemes to compute boundary layer features and obtain reasonable loss predictions.

The relative importance of convection by secondary flows and diffusion by turbulence as mechanisms responsible for mixing in multistage, axial-flow compressors has been investigated by using the ethylene tracer-gas technique and hot-wire anemometry. The tests were conducted at two loading levels in a large, low-speed, four-stage compressor. The experimental results show that considerable cross-passage and spanwise fluid motion can occur and that both secondary flow and turbulent diffusion can play important roles in the mixing process, depending upon location in the compressor and loading level. In the so-called freestream region, turbulent diffusion appeared to be the dominant mixing mechanism. However, near the endwalls and along airfoil surfaces at both loading levels, the convective effects from secondary flow were of the same order of magnitude as, and in some cases greater than, the diffusive effects from turbulence. Calculations of the secondary flowfield and mixing coefficients support the experimental findings.