R. Ajith Kumar

Effect of Corner Cut on the Near Wake Flow Structures of a Square Cylinder

In this paper, results of a flow visualization study on the flow around a square section cylinder with corner chamfering are presented. The corners of the cylinder are chamfered so that the each corner forms a triangle with horizontal (stream-wise) and cross stream (perpendicular to the free stream velocity) dimension ‘b’. Experiments are conducted for b/B0 ratios of 0.05, 0.1, 0.2 and 0.3 where ‘B0’ is the side dimension of the uncut square cylinder. The flow structures, particularly the vortex shedding mode and mechanism around the cylinder with chamfered corners are investigated in order to deduce the effect of corner modifications on the flow. For studies with stationary cylinder (case (a)), the results are taken at Reynolds number values of 1500, 2100 and 2800. For sinusoidally oscillated cylinder case (case (b)), the studies are restricted to Re=2100. To bring out the effect of corner chamfering more clearly, experiments are also conducted with a square cylinder without corner cuts, i.e., with sharp corners. For the case (b), a special mechanism is made to oscillate the cylinders at a desired amplitude and frequency. That is, the cylinder undergoes forced sinusoidal oscillation in case (b). It is found that drag decreases and Strouhal number increases with b/B0 ratio. Quite uniquely, at b/B0=0.2, cross-stream convection of vortices have been observed. Vortex coalescence is observed in almost all cases. Results indicate that corner chamfering brings notable changes in the near-wake flow structures of a square section cylinder. In view of marine structures and building sections with similar geometries, the present results carry considerable practical significance.© 2014 ASME

Numerical Analysis of Effects of Turbine Blade Tip Shape on Secondary Losses

This paper deals with a study of the effect of some turbine blade tip shapes on the secondary flows and the associated aerodynamics. A conventional plain tip shape and a novel squealer tip shape are compared aerodynamically using numerical analysis. The simulations are done using a finite volume-based, general purpose CFD solver, ANSYS FLUENT. The investigation was carried out on a turbine blade cascade consisting of three blades, test blade being the central blade which was modelled. The cascade analysis was done to capture the secondary flows and associated losses. Two cases of tip clearance viz., 0 and 1.5 % of the blade span were considered in this study contributing to the effect of blade tip geometry. The vorticity magnitude at a selected downstream vertical plane was estimated to aerodynamically compare the tip shapes employed in this study. Due to tip clearance, local secondary flows are found to be generated at the blade tip region. Results obtained in this study further indicate that squealer blade tip reduces the secondary flow losses when compared to the conventional plain turbine blade tips. Reduction in secondary flow losses is expected to subdue the effect of heat loads on blade tips. This is perhaps the most prominent practical implication of this key result. The magnitude of vorticity at the blade tip region for squealer tip with 1.5 % tip clearance is 21.25 % less than that for plain tip at the blade tip region for the same tip clearance. This is possibly because of separation of flow and recirculation at the squealer rim which induces weak leakage flows.

CFD Analysis to Investigate the Effect of Vortex Generators on a Transonic Axial Flow Compressor Stage

The performance of the compressor blade is considerably influenced by secondary flow effects, like the cross flow on the end wall as well as corner flow separation between the wall and the blade. Computational Fluid Dynamics (CFD) has been extensively used to analyze the flow through rotating machineries, in general and through axial compressors, in particular. The present work is focused on the studying the effects of Vortex Generator (VG) on test compressor at CSIR National Aerospace Laboratories, Bangalore, India using CFD. The compressor consists of NACA transonic rotor with 21 blades and subsonic stator with 18 vanes. The design pressure ratio is 1.35 at 12930 RPM with a mass flow rate of 22 kg/s. Three configurations of counter rotating VGs were selected for the analysis with 0.25δ, 0.5δ and δ height, where δ was equal to the physical thickness of boundary layer (8mm) at inlet to the compressor rotor [11]. The vortex generators were placed inside the casing at 18 percent of the chord ahead to the leading edge of the rotor. A total of 63 pairs of VGs were incorporated, with three pairs in one blade passage. Among the three configurations, the first configuration has greater impact on the end wall cross flow and flow deflection which resulted in enhanced numerical stall margin of 3.5% from baseline at design speed. The reasons for this numerical stall margin improvement are discussed in detail.Copyright