In this paper, the modifications induced by the presence of an inlet flow non uniformity on the aerodynamic performance of a nozzle vane cascade are experimentally assessed. Tests were carried out in a six vane linear cascade whose profile is typical of a first stage nozzle guide vane of a modern heavy duty GT. An obstruction was located in the wind tunnel inlet section to produce a non uniform flow upstream of the leading edge plane. The cascade was tested in an atmospheric wind tunnel at an inlet Mach number Ma1 = 0.12, with a high turbulence intensity (Tu1 = 9%) and variable obstruction tangential and axial positions, as well as tangential extension. The presented results show that an inlet flow non uniformity influences the stagnation point position when it faces the vane leading edge from the suction side. A relevant increase of both 2D and secondary losses are observed when the non uniformity is aligned to the vane leading edge. When it is instead located in between the passage it does not affect the stagnation point location, in the meanwhile allowing a reduction in the secondary loss.
Aero-Thermal Performance of a Nozzle Vane Cascade With a Generic Non Uniform Inlet Flow Condition: Part I — Influence of Non Uniformity Location
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Perdichizzi, A, Abdeh, H, Barigozzi, G, Henze, M, & Krueckels, J. "Aero-Thermal Performance of a Nozzle Vane Cascade With a Generic Non Uniform Inlet Flow Condition: Part I — Influence of Non Uniformity Location." Proceedings of the ASME Turbo Expo 2016: Turbomachinery Technical Conference and Exposition. Volume 5A: Heat Transfer. Seoul, South Korea. June 13–17, 2016. V05AT13A018. ASME. https://doi.org/10.1115/GT2016-57438
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