Effects of Shock Wave Development on Secondary Flow Behavior in Linear Turbine Cascade at Transonic Condition

Gas turbines widely applied to power generation and aerospace propulsion systems are seriously required to be enhanced in efficiency for the reduction of environmental load. Therefore, the aerodynamic performances of turbine components constituting the gas turbine engine system have to be continuously improved. The energy recovery efficiency from working fluid in turbine component can be enhanced by the increase of turbine blade loading. However, the increase of turbine blade loading inevitably intensifies the secondary flows such as a horseshoe vertex, a passage vortex and a tip leakage vortex, and consequently increases the associated secondary loss. The development of passage vortex is strongly influenced by the pressure distribution on the endwall, especially by the pitchwise pressure gradient in the cascade passage. In addition, a practical high pressure turbine stage is generally driven with high rotational speed under transonic flow conditions where the shock wave appears and strongly influences the pressure distribution on the endwall as well as the loss generation with the increase of exit Mach number. Therefore, it becomes very important to clarify the effects of the development of shock wave on the secondary flow behavior in order to increase the turbine blade loading without the deterioration of efficiency. 

In this study, two-dimensional and three-dimensional transonic flows in the HS1A linear turbine cascade at the design incidence angle were analyzed numerically by using the commercial CFD code STAR-CD ver. 4.18 with the assumption of steady state compressible flow. The HS1A cascade was designed for the midspan section of a high pressure turbine and publishes a lot of experimental data measured at the midspan related to the profile loss and the isentropic Mach number distribution on the blade surface at the transonic conditions. In the present computations, the isentropic exit Mach number was varied in the wide range from the subsonic to the supersonic conditions in order to examine the effects of development of shock wave caused by the increase of exit Mach number on the secondary flow behavior and the associated loss generation. The low-Reynolds type of SST k-w model was used as the turbulence model. The total pressure and the total temperature were applied uniformly at the inlet boundary. The static pressure evaluated from the isentropic exit Mach number was applied uniformly at the outlet boundary. The computed results were compared with the available experimental data to validate the accuracy of the present results and showed good agreements with them.

        The increase of exit Mach number induced the shock from the pressure side of the trailing edge and connected it with the low pressure region near the blade suction surface on the endwall, and consequently produced the shock across the passage and increased its obliqueness. The increase of obliqueness reduced the cross flow on the endwall by moving the low pressure region near the suction surface toward the trailing edge. As a result, the increase of exit Mach number attenuated the passage vortex.