The flow through a transonic compressor cascade shows a very complex structure due to the occurring shock waves. In addition, the interaction of these shock waves with the blade boundary layer inherently leads to a very unsteady flow behavior. The aim of the current investigation is to quantify this behavior and its influence on the cascade performance as well as to describe the occurring transonic flow phenomena in detail. Therefore, an extensive experimental investigation of the flow in a transonic compressor cascade has been conducted within the transonic cascade wind tunnel of DLR Institute of Propulsion Technology at Cologne. In this process, the flow phenomena were thoroughly examined for an inflow Mach number of 1.21. The experiments investigate both the laminar and the turbulent shock wave boundary layer interaction within the blade passage and the resulting unsteady behavior. The experiments show a fluctuation range of the passage shock wave of about 10% chord for both cases, which is directly linked with a change of the inflow angle and of the operating point of the cascade. Thereafter, Reynolds-averaged Navier–Stokes (RANS) simulations have been performed aiming at the verification of the reproducibility of the experimentally examined flow behavior. Here, it is observed that the dominant flow effects are not reproduced by a steady numerical simulation. Therefore, a further unsteady simulation has been carried out to capture the unsteady flow behavior. The results from this simulation show that the fluctuation of the passage shock wave can be reproduced but not in the correct magnitude. This leads to a remaining weak point within the design process of transonic compressor blades because the working range will be overpredicted. The resulting conclusion of this study is that the use of scale-resolving methods such as LES or the application of DNS is necessary to correctly predict unsteadiness of the transonic cascade flow and its impact on the cascade performance.

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