This paper presents a numerical study on blade vibration for the transonic compressor rig at the Technische Universität Darmstadt (TUD), Darmstadt, Germany. The vibration was experimentally observed for the second eigenmode of the rotor blades at nonsynchronous frequencies and is simulated for two rotational speeds using a time-linearized approach. The numerical simulation results are in close agreement with the experiment in both cases. The vibration phenomenon shows similarities to flutter. Numerical simulations and comparison with the experimental observations showed that vibrations occur near the compressor stability limit due to interaction of the blade movement with a pressure fluctuation pattern originating from the tip clearance flow. The tip clearance flow pattern travels in the backward direction, seen from the rotating frame of reference, and causes a forward traveling structural vibration pattern with the same phase difference between blades. When decreasing the rotor tip gap size, the mechanism causing the vibration is alleviated.

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.

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.

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 (BTC). In order to experimentally validate the relation between BTC and aerodynamic damping, the compressor tests were repeated with enlarged BTC 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 BTC and a nonaxisymmetric casing treatment (NASCT). 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 NASCT is validated by means of the unsteady compressor data. By giving a description of all of the 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 NASCT is beneficial for the design of next generation casing treatments for vibration control.

Variable inlet guide vanes enhance the efficiency and stability of modern transonic compressors. The current quest for compact and highly efficient aero-engines requires for higher stage-loading and small axial gaps between adjacent blade rows, increasing interaction between blade rows and introducing further unsteadiness into the flow. Modeling these interactions is relevant to jet engine implementation. Recent advancements in numerical simulation of unsteady flow in multiple blade rows create an additional way to investigate the flow patterns formed by the unsteady interaction. In the case of transonic compressors, the experimental and numerical database is small. The current article will contribute to this database by investigating the flow downstream of an inlet guide vane and the influence this flow has on the passage in a transonic compressor. Unsteady flow phenomena are resolved by piezoresistive wall pressure tappings. Numerical and experimental data show that the interaction of blade rows influences the formation of the tip leakage vortex. Analyzing the vortex structure within the transonic compressor stage it is possible to show how the blade rows interact and how geometric modifications in the VIGV tip gap model influence simulation results.