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Turbine Blade Aerodynamics
Sumanta Acharya
Gazi Mahmood
Louisiana State University CEBA 1419B, Mechanical Engineering Department Baton Rouge, LA 70803
phone: (225) 578-5809 email:
4.3-1 Introduction
The aerodynamics of the flow in a turbine stage (stator/rotor) is rather complex and is still the subject of many ongoing research activities in the gas turbine community. The flow is inherently three dimensional due to the vane/blade passage geometry with features such as twisting of the vane/blade along the span, clearance between the blade tip and the shroud, film cooling holes, and end wall contouring1. The passage flow is characterized by boundary layer effects, secondary flows generated by the passage pressure gradients, and vortical flow structures such as the leading edge horse-shoe vortices, tip-leakage flow vortices, and corner vortices2. The effects of centrifugal-buoyancy, shock- boundary layer interaction, and flow interactions between the stator and rotor rows complicate the passage flow field even further. Along the end walls, the flow structure is strongly three- dimensional with the passage vortex and coolant injection on the hub side and the tip-leakage vortex on the tip side. In the mid- span regions located away from the passage walls and outside the viscous shear layer, the radial flow is almost negligible and the flow is effectively two dimensional. The fluid dynamics in this region can then be based on two dimensional planar cascade flow studies without any significant loss of information. The three dimensional complex flow structures near the hub endwall region and in the blade tip-shroud clearance have been simulated in annular vane/blade passages with and without rotating blade row3. Studies of the complex end-wall flows have also been performed in stationary cascades with three dimensional airfoil shapes4. The qualitative features of the passage flows, which comprise mainly of the passage crossflow (flow from the pressure side of vane/blade to suction side of adjacent vane/blade) and vortical flows induced by the leading edge, the corners, and the injected coolant flows have been studied in detail in stationary cascades and are considered to be similar in both stationary and rotating blade rows. The primary difference in the secondary flow structure between the blade passage and vane passage is that the vortical flows and cross flows in the blade passage are stronger because of higher turning of the flows along the blade passage. Secondary flows are the major source of aerodynamic losses, which account for 35%-40% of all losses5 and thermal loading in the turbine passage, and thus require special considerations by the turbine designers.
The primary objectives of this chapter are to present and analyze the features of the flow field in the turbine vane/blade passage near the hub endwall and mid-span locations of the blade. Toward this effort, reported measurements and computations of pressure, velocity distributions, flow turning angles, turbulence intensity, and vorticity distributions in the cascade test section are presented. Recent efforts to reduce the secondary flows by structural modifications in the passage are discussed. In this chapter, basic fluid dynamic principles and mathematical models of the flow in the passage are not discussed, and the reader is referred to notes 1, 2, and 6 for additional details6. Also details on the aerodynamic design methodology for the vane/blade passage are not presented.

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