Tubercles are modifications to the leading edge of an airfoil in the form of blunt wave-like serrations. Several studies on the effect of tubercles on isolated airfoils have shown a beneficial effect in the post-stall regime, as reduced drag and increased lift, leading to a delay of stall. The prospect of delaying stall is particularly attractive to designers of axial compressors in gas turbines, as this leads to designs with higher loading and therefore higher pressure rise with fewer number of stages. In the present study, experiments were performed on a cascade of airfoils with NACA 65209 profile with different tubercle geometries. The measurements were made over an exit plane using a five-hole probe to compare the cascade performance parameters. Additionally, hot-wire measurements were taken near the blade surface to understand the nature of the flow in the region close to the tubercles. Oil-flow visualization on the cascade end wall reveal the flow through the passage of blades with and without tubercles. For the cascade considered, the estimated stall angle for the best performing set of blades is found to increase up to 8.6° from that of the unmodified blade of 6.0°. Application of such structures in axial compressor blades may well lead to suppression of stall in axial compressors and extend the operating range. NOMENCLATURE l blade chord s blade pitch ξ stagger angle i incidence angle
In the present study, the effectiveness of passive structures called tubercles on an axial compressor blade row is studied experimentally. Tubercles are the modifications to the leading edge of an airfoil in the form of blunt wave-like serrations. Although several studies on the effect of tubercles on isolated blades are available in literature, detailed study of their effect on a cascade of blades, such as in the case of an axial flow turbo-machine is lacking. Such an application in an axial compressor will result in a significant increase in the stall margin. Presently, experiments have been performed on a linear compressor cascade with a blade height of 0.15 m and mean chord of 0.06 m, on a NACA 65209 airfoil profile. The plain and modified blades are fabricated using rapid prototyping to ensure conformity to the required geometry. The cascade is designed in such a way that the incidence (angle of attack) and the stagger can be changed easily. The measurements are taken at the exit plane using a five-hole Pitot probe to obtain three-components of velocity and static pressure data over fine measurement grids. The effect is determined in terms of lift and drag coefficients, lift-to-drag ratio and total pressure loss coefficient. Experiments have been carried out for different pitch and amplitude (serration depth) of tubercles to understand their effect. The stall incidence angle for the best performing blade is found to increase up to 8.6° from that of the unmodified blade of 6.0°. Application of such structures in axial compressor blades may well be adequate to prevent stalling in axial compressors over a wide operating range.
The intakes of modern aircraft are subjected to ever-increasing demands in their performance. Particularly, they are expected to carry out diffusion with the highest isentropic efficiency while subjected to aggressive geometry requirements arising from stealth considerations. To avoid a penalty in engine performance, the flow through intake needs to be controlled using various methods of flow control. In this study, a serpentine intake is studied experimentally and its performance compared with and without boundary layer suction. The performance parameters used are nondimensional total pressure loss coefficient and standard total pressure distortion descriptors. The effect is observed on surface pressure distributions, and inferences are made regarding separation location and extent. A detailed measurement at the exit plane shows flow structures that draw attention to secondary flows within the duct. Suction is applied at three different locations, spanning different number of ports along each location, comprising of ten unique configurations. The mass flow rate of suction employed ranges from 1.1% to 6.7% of mass flow rate at the inlet of the intake. The effect is seen on exit total pressure recovery as well as circumferential and radial distortion parameters. This is examined in the context of the location of the suction ports and amount of suction mass flow, by the deviation in surface pressure distributions, as well as the separation characteristics from the baseline case. The results show that applying suction far upstream of the separation point together with a modest amount of suction downstream results in the best performance.
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