## 1 Introduction
Although the RANS approach overestimated the size of the separation bubble, it captured the overall pattern accurately. The misprediction could be attributed mainly to the unsteady characteristics of corner separation [15; 16]. In the analysis of transonic/supersonic compressor cascades, it is crucial to accurately capture the interaction between shock waves and boundary layers. This interaction significantly affects the performance estimation of turbomachinery [17; 18]. Therefore, establishing a correlation between grid resolution, turbulence model, and solver settings is necessary to achieve a precise representation of the fluid flow and ensure proper validation against experimental data [19]. A study by [PERSON] et al. [20] demonstrated the substantial impact of the turbulence model on performance prediction. This influence was primarily attributed to the high sensitivity of shockwave patterns to the development of the boundary layer. In the field of supersonic cascades, [PERSON] et al. [21] developed a one-dimensional analytical model for predicting shock losses, aiming to facilitate the rapid evaluation and reduction in these losses in the optimization design of a supersonic compressor cascade. Their model was applied to two supersonic cascades, resulting in a significant reduction of 20% and 16% in the total pressure loss coefficient. Furthermore, [PERSON] et al. [22] conducted research to determine the primary source of shock losses. By developing a shock loss mechanism and validating the results, it was identified that the main source of loss in the supersonic cascade is the terminal passage shock loss. Other related studies focusing on the development of shock loss prediction methods are found in [23; 24; 25], and most of them were developed based on experimental empirical correlations. These studies contribute to the understanding of shock losses in supersonic cascades and provide insights into the development of analytical models and optimization approaches to mitigate these losses, ultimately enhancing the performance of supersonic compressor cascades. To control the flow field inside a cascade wind tunnel and establish a procedure for this approach, [PERSON] et al. [8] developed a numerical method capable of simulating complex configurations involving movable mechanisms. Using that method, steady and unsteady simulations of a transonic cascade wind tunnel were performed. Their study investigated the effects of wall suction, throttle opening angle, and tailboard angle in the case of steady analysis, as well as blade oscillation in the case of unsteady analysis. The findings revealed that the differences between the ideal configuration and the wind tunnel configuration become more prominent as the pressure amplitude increases downstream of the lower blades. Furthermore, this research study highlights the significance of various parameters on the flow characteristics and contributes to understanding the control and assessment of flow fields within cascade wind tunnels using numerical simulations. Similar studies focusing on advanced analytical techniques and computational tools to address complex problems are presented in [26; 27; 28]. The current work examines the effects on the flow structure of an asymmetric transonic linear cascade in contrast with single-passage mapping, with the study being performed for a mixed flow/centrifugal compressor vaned diffuser. The primary focus of this paper is centered around the development of an axisymmetric test rig designed to facilitate the visualization of shock waves on diffuser vanes. Managing the interaction between shock waves and the boundary layer is crucial for minimizing drag and shock losses. The presence of such an interaction introduces aerodynamic instabilities that significantly impact the overall performance and operational capabilities of turbomachinery. Using an asymmetric cascade configuration to scrutinize flow patterns and the structure of shock waves has proven to be fruitful, leading to reduced computational and time loads.