3D Transonic Flow in a Compressor Cascade With Shock-Induced Corner Stall

Abstract

An experimental and numerical study of the transonic flow through a linear compressor cascade with endwalls was conducted. The cascade with a low aspect ratio of 1.34 was tested at an inlet Mach number of 1.09 and a Reynolds number of 1.9×106. Detailed flow visualizations on the surfaces and 5-hole probe measurements inside the blading and in the wake region showed clearly a 3-dimensional boundary layer separation on the blade surface and the sidewall, and a severe corner stall induced by a strong 3D shock system at blade passage entrance. The experimental data has been used to validate and improve the 3D Navier-Stokes code TRACE. Results showed an excellent resolution of the complex flow field. Surface pressure distributions on the entire blade surface and the endwalls, flow angle and total pressure contours within the blade passage and the wake are compared with the experimental results. An analysis of the secondary flow of this highly staggered cascade did not show the classical corner vortex. Instead, a severe flow deviation and partly reverse flow near the walls is seen. The flow solver helped to identify a weak ring vortex that originates from the passage sidewall. Surface oil flow pictures on the blade contour and the sidewall are in qualitatively good agreement to numerical surface streaklines. A considerable improvement of the numerical results could be achieved by a gradual grid refinement especially in the corner region and by successive code development.

Organized Streamwise Vorticity on Convex Surfaces With Particular Reference to Turbine Blades

“…A theory giving expected spanwise wavelengths for a circular cylinder was developed by Kestin and Wood 9 in 1970. "Streaky structures" and streamwise vortices have been observed and predicted on flat plates 10 , on compressor blades 11,12 and on the convex surface of a supercritical airfoil, following an initial concave surface exhibiting Görtler vorticity 13 . In a study of plane shear layers, Lasheras et al 14 have concluded that these are unstable to three-dimensional perturbations in the upstream conditions.…”

Experimental investigation on a high subsonic compressor cascade flow

“…Through the years, to improve the work performance of highly loaded axial compressor, the structure of corner separation has been studied worldly in the aeronautical research field. [1][2][3][4][5][6][7][8][9][10][11] Compressor cascade is a simplified model of real axial compressor stator blade and has been widely applied to researching the mechanism and control of corner separation, as well as the compressor tip clearance flow structures and effects of threedimensional compressor blades. 12,13 Active flow control methods, such as boundary layer suction, [14][15][16] flow blowing, 17,18 plasma flow control 19,21 etc, have been applied to the control of compressor cascade corner separations over years.…”

“…To control the corner separation effectively, the design of both active and passive methods should be based on the in-depth understanding of flow structures in the compressor cascade flow passage and various efforts have been spared. [1][2][3][4][5][6][7][8][9][10][11][14][15][16][17][18][19][20][21][22][23][24][25][26][27] In low speed flow, with inlet Mach number under 0.3, detailed researches on the flow structures in the compressor cascade flow passage have been launched. 3,[7][8][9][10] When inlet Mach number increases to be bigger than 0.3, the flow becomes more unsteady and compressible.…”

“…Under the real work conditions, compressor blades are more likely to encounter high subsonic fluid, and research groups of German Aerospace Center Deutsches Zentrum fu¨r Luft-und Raumfahrt (DLR), [22][23][24][25][26] Tiedemann 17 and Zhang et al 20 have carried out some trial study to control the corner separation with the inlet Mach number around 0.7. While the research on control separation in high-speed flow is under way, compressor cascade flow structure has been studied, too, 5,11,16,17,20,[22][23][24][25][26] but is still in lack of detailed investigation compared with that under low-speed flow circumstances. So it is meaningful to provide detailed research of compressor cascade flow structure as the basis for better understanding and controlling of corner separations in high subsonic flow.…”

The Influence of Streamwise Vorticity on Transition in the Presence of Sweep

“…Previous investigations have revealed streamwise vortices and "streaky structures" on flat plates 1 and on the suction surface of compressor blades 2,3 . Turbine blade designers are quite familiar with the phenomenon of Görtler vorticity; this is thought to occur predominantly on the concave pressure surfaces of turbine blades.…”