Abstract
Modern aeronautic fans are characterised by a transonic flow regime near the blade tip. Transonic cascades enable higher pressure ratios by a complex system of shockwaves arising across the blade passage, which has to be correctly reproduced in order to predict the performance and the operative range. In this paper, we present an accurate two-dimensional numerical modelling of the ARL-SL19 transonic compressor cascade. A large series of data from experimental tests in supersonic wind tunnel facilities has been used to validate a computational fluid dynamic model, in which the choice of turbulence closure resulted critical for an accurate reproduction of shockwave-boundary layer interaction. The model has been subsequently employed to carry out a parametric study in order to assess the influence of main flow variables (inlet Mach number, static pressure ratio) and geometric parameters (solidity) on the shockwave pattern and exit status. The main objectives of the present work are to perform a parametric study for investigating the effects of the abovementioned variables on the cascade performance, in terms of total-pressure loss coefficient, and on the shockwave pattern and to provide a quite large series of data useful for a preliminary design of a transonic compressor rotor section. After deriving the relation between inlet and exit quantities, peculiar to transonic compressors, exit Mach number, mean exit flow angle and total-pressure loss coefficient have been examined for a variety of boundary conditions and parametrically linked to inlet variables. Flow visualisation has been used to describe the shock-wave pattern as a function of the static pressure ratio. Finally, the influence of cascade solidity has been examined, showing a potential reduction of total-pressure loss coefficient by employing a higher solidity, due to a significant modification of shockwave system across the cascade.
Highlights
The development of modern aircraft compressors aims to increase pressure ratios and to reduce weights at once [1,2]
The cascade analysed in this work is known as Aerospace Research Laboratories (ARL)-SL19 supersonic compressor cascade and derives from a rotor near-tip section of a transonic axial compressor described in [13,14]
The present work provides a detailed study of the two-dimensional flow in a linear supersonic compressor cascade for different operating conditions, which may be of support to the preliminary design of a transonic/supersonic compressor cascade, the corresponding rotor blade section, or the stage of a supersonic compressor or fan
Summary
The development of modern aircraft compressors aims to increase pressure ratios and to reduce weights at once [1,2]. A two-dimensional Navier–Stokes solver was validated using experimental data obtained in a supersonic cascade facility in order to investigate the influence of inlet Mach number and axial velocity-density ratio (AVDR) on blade performance. The main independent variables are flow quantities, such as the inlet Mach number, the static pressure ratio, and the AVDR, or geometric parameters, such as the cascade solidity. All of these variables strongly influence the cascade performance, in terms of total-pressure loss coefficient, and the exit flow field, in terms of exit Mach number and mean exit flow angle. In the remainder of the paper, the validation of the numerical model is thoroughly reported, followed by the outcome of the parametric study, which can be useful for the preliminary design of a transonic compressor rotor blade section
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