Due to its high influence on the flow behavior, the interaction of shock waves with boundary layers is subject of numerous research projects since decades. The dynamics emerging from the shock as well as from initiated instabilities in the boundary layer may trigger structural excitation and drag rise. Particularly in turbomachinery applications, harmonic aerodynamic loads can potentially cause hazardous oscillations of the blade and, therefore, it is of major interest during the aerodynamic design process. The importance of this subject rises with the promoted development on geared turbofans. Due to the high spool speed of the low pressure turbine, the flow on the turbine blade suction surface becomes transonic particularly in blade tip regions. Of particular relevance in this context is the interaction of the shock wave generated at the trailing edge with the boundary layer on the suction surface of the neighboring blade, which is a well-known phenomenon for supersonic exit Mach numbers.

The results presented in the paper are based on experimental investigations at a transonic generic low pressure turbine profile but at subsonic exit Mach numbers. Here, the flow on the suction side reaches a maximum isentropic Mach number of approximately 1.2 and features a large separation bubble in a transonic flow regime. In this case the terminating shock wave on the suction side is not impinging from the neighboring blade but appears far downstream of the trailing edge. Thereby, the interaction between the shock wave and the separated flow leads to a reattachment of the flow. The experiments were carried out at low density conditions in the High Speed Cascade Wind Tunnel at the Institute of Jet Propulsion of the Bundeswehr University Munich. The test facility enables an independent adjustment of the Mach and Reynolds number as well as the turbulence level. Therefore, this test facility is highly suitable to investigate flow phenomena of turbine profiles in the transonic flow regime at engine relevant conditions.

Based on measurements of the static pressure at the profile surface and the Surface Hot Film Sensors the boundary layer conditions for different operating points will be discussed in the paper. The measuring of the quasi wall shear stresses by the Hot Film Sensors provides a revealing insight into the flow separation, its extent, and the transitional behavior of the boundary layer. Furthermore, the influence of the Reynolds and Mach number from different operating points is discussed in detail. It is moreover demonstrated, that the final normal shock strongly dominates the transition of the boundary layer flow, which enables the reattachment of the pervious separated flow. The measurements are supplemented by Schlieren images recorded by a high speed camera at 19.2 kHz. From this recordings a high unsteadiness of the shock waves system is observable, which can clearly related to its interaction with the separated boundary layer. A statistical analysis of the images is performed by processing the time signal of the intensity of each pixel. This reveals characteristic frequencies of the shock waves and the visible shear layers. An analogue analysis of the time signal from the Surface Hot Film Sensors is showing the instabilities influencing the boundary layer which can be furthermore linked to the unsteadiness appearing in the main flow.


Dynamics of Shock Waves Interacting With Laminar Separated Transonic Turbine Flow Investigated by High-Speed Schlieren and Surface Hot-Film Sensors

Division / Sponsor:
International Gas Turbine Institute (IGTI)

Paper Proceedings: https://asmedigitalcollection.asme.org/GT/GT2020/volume/84058

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