In the recent past, experimental studies have shown some advantages of blade lean and sweep in axial compressors. As most of the experimental results are combined with other features, it is difficult to determine the effect of individual parameters on the performance of the compressor. The present numerical studies are aimed at understanding the performance and three-dimensional flow pattern within and at the exit of swept and unswept rotors. Three rotors, namely, unswept, 20° forward swept, and 20° backward swept rotors, are analysed with a specific intention of understanding the three-dimensional flow pattern within the rotors and also the pattern of the blade boundary layer flow. The analysis was done using a fully three-dimensional viscous CFD code CFX-5. Results indicated a reduction in pressure rise with sweep. Backward sweep adversely affects the stall margin. Forward sweep changes the streamline pattern in such a way that the suction surface streamlines are deflected towards the hub and the pressure surface streamlines are deflected towards the casing. An opposite behaviour is observed in the backward swept rotors. High axial velocities reduce the secondary losses near the hub, resulting in a high pressure rise in forward swept rotor.

The present study comprises steady state, two-dimensional computational investigations performed on NACA 0012 airfoil to analyze the effect of Gurney flap (GF) on airfoil aerodynamics using k-ε RNG turbulence model of FLUENT. Airfoil with GF is analyzed for six different heights from 0.5% to 4% of the chord length, seven positions from 0% to 20% of the chord length from the trailing edge, and seven mounting angles from 30° to 120° with the chord. Computed values of lift and drag coefficients with angle of attack are compared with experimental values and good agreement is found at low angles of attack. In addition static pressure distribution on the airfoil surface and pathlines and turbulence intensities near the trailing edge are present. From the computational investigation, it is recommended that Gurney flaps with a height of 1.5% chord be installed perpendicular to chord and as close to the trailing edge as possible to obtain maximum lift enhancement with minimum drag penalty.

The present paper reports experimental investigations on the effect of diffuser vane height and position on the performance of a low-speed centrifugal compressor. The diffuser vane height is systematically varied from 0.2 to 0.9 times the diffuser width. In addition, the effect of vane position is examined by fixing the partial vanes to the hub, shroud, or hub and shroud. The compressor performance is determined with these vanes in the vane and low solidity vane diffuser configurations. It is found that there is an optimum height for the diffuser vane height. In the present investigation, it is found to be 0.3 times the diffuser width. The effect of position of the partial vanes on the hub or shroud on the compressor performance is found to be negligible. However, when partial vanes are fixed on the hub and shroud staggered at half spacing, the compressor performance is improved substantially.

Interaction between rotating impeller and stationary diffuser in a centrifugal compressor is of practical importance in evaluating system performance. The present study aims at investigating how the interaction influences the unsteady diffuser performance and understanding the physical phenomena in the centrifugal compressor. A computational fluid dynamics (CFD) method has been applied to predict the flow field in the compressor, which has a conventional vaned diffuser (VD) and a low solidity vaned diffuser (LSVD). The radial gaps between impeller and diffuser and different flow coefficients are varied. The results obtained show that the major parameter that influences the unsteady variation of diffuser performance is due to the circumferential variation of the flow angle at the diffuser vane leading edge. The physical phenomena behind the pressure recovery variation are identified as the unsteady vortex shedding and the associated energy losses. The vortex core region as well as the shedding of vortices from the diffuser vane are triggered by the variation in the diffuser vane loading, which in turn is influenced by the circumferential variation of the impeller wake region. There is little unsteady variation of flow angle in the span-wise direction. This indicates that the steady state performance characteristics are related to the span-wise variation of flow angle, while the unsteady characteristics are contributed by the circumferential variation of flow angle. At design conditions, dominant frequency components of pressure fluctuation are all periodic and at near stall, these are aperiodic.

A computational study has been conducted to analyse the performance of a centrifugal compressor with different types of diffusers under various levels of impeller—diffuser interactions. Vaneless (VLD), vaned (VD), low solidity vaned (LSVD), and partial vaned diffusers (PVD) are used for this purpose. The study is carried out using commercial software ANSYS CFX. The interaction level is varied by varying the radial gap between the impeller and diffuser by keeping the diffuser vane at three different radial locations. Numerical simulations have been conducted for four different flow coefficients. At design flow coefficient maximum efficiency occurs when the leading edge is at R3 (ratio of radius of the diffuser leading edge to the impeller tip radius) = 1.10 for all vane-type diffuser configurations. At below design flow coefficient higher stage efficiency occurs when the diffuser vanes are kept far away ( R3 = 1.15) and at above design flow coefficient R3 = 1.05 gives better efficiency. The highest diffuser pressure recovery coefficient ( Cp) is observed for VD at design flow coefficient. For VLD, the Cp value increases with flow coefficient. In the case of VD and LSVD configurations the exit flow from the impeller is disturbed when the diffuser vanes are closer, and these disturbances are more evident in the last 10 per cent of the impeller flow. In the case of the impeller with PVD the interaction effects are minimum.

The blade-to-blade variation of relative stagnation pressure losses in the tip region inside the rotor of a single-stage, axial-flow compressor is presented and interpreted in this paper. The losses are measured at two flow coefficients (one at the design point and the other at the near peak pressure rise point) to discern the effect of blade loading on the end-wall losses. The tip clearance losses are found to increase with an increase in the pressure rise coefficient. The losses away from the tip region and near the hub regions are measured downstream. The losses are integrated and interpreted in this paper.

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This paper reports the measurement of the relative flow in the rotor blade passages of an axial-flow compressor and an axial-flow inducer using conventional probes such as five-hole, disk, and spherical pitot-static probes. The five-hole probe provides an inexpensive, yet accurate, method of deriving the three-dimensional flow field in a rotor; the disc probe is suitable for the blade boundary-layer measurement; and the pitot-static probe is useful for obtaining the static and stagnation pressures and the total velocity across the rotor passage. Typical data obtained from each of these probes demonstrate the complex nature of the turbomachinery rotor flow. A detailed discussion of various sources of probe errors and methods of estimating these errors is also included.