Modern aeronautic fans are characterised by a transonic flow regime near the blade tip.\nTransonic cascades enable higher pressure ratios by a complex system of shockwaves arising across\nthe blade passage, which has to be correctly reproduced in order to predict the performance and the\noperative range. In this paper, we present an accurate two-dimensional numerical modelling of the\nARL-SL19 transonic compressor cascade. A large series of data from experimental tests in supersonic\nwind tunnel facilities has been used to validate a computational fluid dynamic model, in which the\nchoice of turbulence closure resulted critical for an accurate reproduction of shockwave-boundary\nlayer interaction. The model has been subsequently employed to carry out a parametric study in\norder to assess the influence of main flow variables (inlet Mach number, static pressure ratio) and\ngeometric parameters (solidity) on the shockwave pattern and exit status. The main objectives of the\npresent work are to perform a parametric study for investigating the effects of the abovementioned\nvariables on the cascade performance, in terms of total-pressure loss coefficient, and on the shockwave\npattern and to provide a quite large series of data useful for a preliminary design of a transonic\ncompressor rotor section. After deriving the relation between inlet and exit quantities, peculiar to\ntransonic compressors, exit Mach number, mean exit flow angle and total-pressure loss coefficient\nhave been examined for a variety of boundary conditions and parametrically linked to inlet variables.\nFlow visualisation has been used to describe the shock-wave pattern as a function of the static pressure\nratio. Finally, the influence of cascade solidity has been examined, showing a potential reduction\nof total-pressure loss coefficient by employing a higher solidity, due to a significant modification of\nshockwave system across the cascade.
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