The increase of the thrust/weight ratio of aircraft engines is extremely restricted by different 3-D flow loss mechanisms. One of them is the corner separation that can form at the junction between a blade suction side and a hub or shroud. In this paper, in order to further investigate the turbulent characteristics of corner separation, large-eddy simulation (LES) is conducted on a compressor cascade configuration using NACA65 blade profiles (chord based Reynolds number: 3.82 × 10⁵), in comparison with the previous obtained experimental data. Using the shear-improved Smagorinsky model as subgrid-scale model, the LES gives a good description of the mean aerodynamics of the corner separation, especially for the blade surface static pressure coefficient and the total pressure losses. The turbulent dynamics is then analyzed in detail, in consideration of the turbulent structures, the one-point velocity spectra, and the turbulence anisotropy. Within the recirculation region, the energy appears to concentrate around the largest turbulent eddies, with fairly isotropic characteristics. Concerning the dynamics, an aperiodic shedding of hairpin vortices seems to induce an unsteadiness of the separation envelope.

This paper considers the inherent unsteady behavior of the three-dimensional (3D) separation in the corner region of a subsonic linear compressor cascade equipped of 13 NACA 65-009 profile blades. Detailed experimental measurements were carried out at different sections in spanwise direction achieving, simultaneously, unsteady wall pressure signals on the surface of the blade and velocity fields by time-resolved particle image velocimetry (PIV) measurements. Two configurations of the cascade were investigated with an incidence of 4 deg and 7 deg, both at $Re=3.8×105$ and Ma = 0.12 at the inlet of the facility. The intermittent switch between two statistical preferred sizes of separation, large, and almost suppressed, is called bimodal behavior. The present PIV measurements provide, for the first time, time-resolved flow visualizations of the separation switch with an extended field of view covering the entire blade section. Random large structures of the incoming boundary layer are found to destabilize the separation boundary. The recirculation region, therefore, enlarges when these high vorticity perturbations blend with larger eddies situated in the aft part of the blade. Such a massive detached region persists until its main constituting vortex suddenly breaks down and the separation almost completely vanishes. The increase of the blockage during the separation growth phase appears to be responsible for this mechanism. Consequently, the proper orthogonal decomposition (POD) analysis is carried out to decompose the flow modes and to contribute to clarify the underlying cause-effect relations, which predominate the dynamics of the present flow scenario.

This paper describes the investigations performed to better understand unsteady flows that develop in a three-stage high-pressure compressor. More specifically, this study focuses on rotor-stator interactions and tip leakage flow effects on overall performance and aerodynamic stability. The investigation method is based on three-dimensional unsteady RANS simulations, considering the natural spatial periodicity of the compressor. Indeed, all information related to rotor-stator interactions can be computed. A comparison is first done with experimental measurements to outline the capacity of the numerical method to predict overall performance and unsteady flows. The results show that the simulation correctly estimates most flow features in the multistage compressor. Then numerical data obtained for three configurations of the same compressor are analyzed and compared. Configurations 1 and 2 consider two sets of tip clearance dimensions and a casing treatment based on a honeycomb design is applied for configuration 3. Detailed investigations of the flow at the same operating line show that the tip leakage flow is responsible for the loss of stability in the last stage. An increase by 30% of the tip clearance dimension dramatically reduces the stable operating range (by 40% with respect to the standard configuration). A modal analysis shows that the stall process in this case involves the perturbation of the flow in the last rotor by upstream stator wakes, leading to the development of a rotating instability. The control device designed and investigated in this study allows for reducing the sensitivity of the compressor to tip leakage flow by recovering the initial stable operating range.

A novel test facility for transonic fans has been constructed and commissioned at École Centrale de Lyon (ECL) within the Project PHARE-2. The facility is instrumented with multi-physical measurement systems to deeply investigate aeroelastic and aeroacoustic phenomena. To enable long-term fundamental research, the composite material open-test-case rotor ECL5 is currently under development, which shall be established as a new reference case for method development. Within the present publication, the project objectives, the current rotor design, ECL5v2, and details on the test facility are presented. With the goal to establish future academic collaborations, the authors aim to initiate a discussion in the research community in an early stage of the test case development to receive constructive feedback on the design and research approach.

The application of composite fans enables disruptive design possibilities but increases sensitivity to multi-physical resonance between aerodynamic, structure dynamic and acoustic phenomena. As a result, aeroelastic problems increasingly set the stability limit. Test cases of representative geometries without industrial restrictions are a key element of an open scientific culture but are currently non-existent in the turbomachinery community. In order to provide a multi-physical validation benchmark representative of near-future UHBR fan concepts, the open-test-case fan stage ECL5 was developed at Ecole Centrale de Lyon. The design intention was to develop a geometry with high efficiency and a wide stability range that can be realized using carbon fibre composites. This publication aims to introduce the final test case, which is currently fabricated and will be experimentally tested. The fan blades are composed of a laminate made of unidirectional carbon fibres and epoxy composite plies. Their structural properties and the ply orientations are presented. To characterize the test case, details are given on the aerodynamic design of the whole stage, structure dynamics of the fan and aeroelastic stability of the fan. These are obtained with a state-of-art industrial design process: static and modal FEM, RANS and LRANS simulations. Aerodynamic analysis focuses on performance and shows critical flow structures such as tip leakage flow, radial flow migration and flow separations. Mechanical modes of the fan are described and discussed in the context of aeroelastic interactions. Their frequency distribution is validated in terms of resonance risk with respect to synchronous vibration. The aeroelastic stability of the fan is evaluated at representative operating points with a systematic approach. Potential instabilities are observed far from the operating line and do not compromise experimental campaigns.