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.

The objective of this research was to use vortex generators to reduce total pressure distortion (i.e., total pressure nonuniformity) and improve total pressure recovery within a curved subsonic diffuser. In this study more than 20 configurations of both co-and counter-rotating arrays of vortex generators were tested in a diffusing S duct. Surface static pressure, surface flow visualization, and exit plane total pressure and transverse velocity data were acquired. The aerodynamic performance of each configuration was assessed by calculating total pressure recovery and spatial distortion elements. The best configuration tested reduced distortion by more than 50% while improving total pressure recovery by 0.5%. The results indicate that the mechanism responsible for improved aerodynamic performance is not boundary-layer re-energization from shed axial vortices but rather the suppression of detrimental secondary flows by redirecting the flow. Nomenclaturegenerator height / = vortex generator lateral spacing N = number of vortex generators p = static pressure po = total pressure JJQ = average total pressure r = radial coordinate s = distance along duct centerline jt, v, z = Cartesian coordinates Apc/p = circumferential distortion intensity element Apr/p = radial distortion intensity element 8 = boundary-layer thickness, 95% 9~ = circumferential distortion angular extent element = polar coordinate Subscripts face = area-weighted face average low = average over 0 for constant r in regions where p 0 is less than po,nn g ; see Fig. 3 ref = reference conditions at s/D\ = -0.5 ring = average over 0 for constant r\ see Fig. 3 1,2 = inlet and exit of duct at s/Di = 0 and 5.23

No abstract

No abstract