Application of nonaxisymmetric casing treatments (CTs) can extend the operating range of a transonic compressor significantly. Recent CT designs have proven successful at achieving operating range extension without efficiency loss under design conditions. Two different CT designs were investigated on a high-speed one and a half stage test rig using extensive instrumentation. The stage setup is representative of the front stage of a modern high-pressure compressor. Results of particle image velocimetry (PIV) measurements taken in the blade tip region underneath the CT show a significantly modified flow structure compared to the smooth casing reference case. Blockage zone, secondary flow, and shock structures are affected by the CT, especially in highly throttled operating conditions. The stall inception process of the system with axial slots shows unexpected behavior, with modal activities that are not observed without CT. These activities are resolved using unsteady wall pressure (WP) and hot wire measurements.

This paper investigates the acoustically induced rotor blade vibration that occurred in a state-of-the-art 1.5-stage transonic research compressor. The compressor was designed with the unconventional goal to encounter self-excited blade vibration within its regular operating domain. Despite the design target to have the rotor blades reach negative aerodamping in the near stall region for high speeds and open inlet guide vane, no vibration occurred in that area prior to the onset of rotating stall. Self-excited vibrations were finally initiated when the compressor was operated at part speed with fully open inlet guide vane along nominal and low operating line. The mechanism of the fluid–structure interaction behind the self-excited vibration is identified by means of unsteady compressor instrumentation data. Experimental findings point toward an acoustic resonance originating from separated flow in the variable inlet guide vanes (VIGV). A detailed investigation based on highly resolved wall-pressure data confirms this conclusion. This paper documents the spread in aerodynamic damping calculated by various partners with their respective aeroelastic tools for a single geometry and speed line. This significant spread proves the need for calibration of aeroelastic tools to reliably predict blade vibration. This paper contains a concise categorization of flow-induced blade vibration and defines criteria to quickly distinguish the different types of blade vibration. It further gives a detailed description of a novel test compressor and thoroughly investigates the encountered rotor blade vibration.

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Inflow distortions in the compression system of a jet engine are becoming increasingly important for research focus. The investigation of the emergence of a distortion, its interaction with the rotor and the resulting impact on the rotor flow is challenging. In this work a separation in the inflow of a transonic compressor was created and the impact on stage aerodynamics investigated. The separation resulted in a total pressure distortion close to the casing within a sector of 120°. Effects were studied both numerically and experimentally in a joint collaboration project. The numerical model consisted of the full rotor-stator compressor stage, the inlet duct and the distortion generator upstream of the stage. This enables both an accurate validation of the numerical results and contributes to a deeper understanding of the flow. The results of both the numerical and experimental studies were in good agreement. The rotor is locally throttled by the inlet separation, resulting in the formation of an additional loss core at the stability limit due to a local aerodynamic overload. Considering classic distortion descriptors like the DC60, it is shown that they are not able to adequately assess the impact of a strong, but small distortion close to the tip of the rotor. The data can be considered as test case for future numerical models as well as for the validation of new analytical models. Furthermore, the results of this study reveal effects in both experimental and numerical studies that would not be realized if only a model of the separation was analyzed.

This paper investigates the vibrations that occurred on the blisk rotor of a 1.5-stage transonic research compressor designed for aerodynamic performance validation and tested in various configurations at Technische Universität Darmstadt. During the experimental test campaign self-excited blade vibrations were found near the aerodynamic stability limit of the compressor. The vibration was identified as flutter of the first torsion mode and occurred at design speed as well as in the part-speed region. Numerical investigations of the flutter event at design speed confirmed negative aerodynamic damping for the first torsion mode, but showed a strong dependency of aerodynamic damping on blade tip clearance. In order to experimentally validate the relation between blade tip clearance and aerodynamic damping, the compressor tests were repeated with enlarged blade tip clearance for which stability of the torsion mode was predicted. During this second experimental campaign, strong vibrations of a different mode limited compressor operation. An investigation of this second type of vibration found rotating instabilities to be the source of the vibration. The rotating instabilities first occur as an aerodynamic phenomenon and then develop into self-excited vibration of critical amplitude. In a third experimental campaign, the same compressor was tested with reference blade tip clearance and a non-axisymmetric casing treatment. Performance evaluation of this configuration repeatedly showed a significant gain in operating range and pressure ratio. The gain in operating range means that the casing treatment successfully suppresses the previously encountered flutter onset. The aeroelastic potential of the non-axisymmetric casing treatment is validated by means of the unsteady compressor data. By giving a description of all of above configurations and the corresponding vibratory behavior, this paper contains a comprehensive summary of the different types of blade vibration encountered with a single transonic compressor rotor. By investigating the mechanisms behind the vibrations, this paper contributes to the understanding of flow induced blade vibration. It also gives evidence to the dominant role of the tip clearance vortex in the fluid-structure-interaction of tip critical transonic compressors. The aeroelastic evaluation of the non-axisymmetric casing treatment is beneficial for the design of next generation casing treatments for vibration control.