Stage Transonic Axial Flow Compressor Research Articles

Geometry scaling technique and application to aerodynamic redesign of multi-stage transonic axial-flow compressors

Geometry scaling technique, as one of the essential steps in aerodynamic redesign of multi-stage compressor, directly determines whether the new scaled geometry meets the required performance or not. It is also one of the effective measures in unsteady flow simulation of multi-stage compressor to reduce greatly the computational cost by scaling the computational domain from full-annulus passages to partial-annulus passages. The common critical issue arisen from both practices is the geometry dissimilarity in which some certain difference exists between the original and scaled geometries due to a diversity of redesign requirements or model simplifications. However, the extent of how the geometry dissimilarity influences the aerodynamic performance of multi-stage transonic compressor is unclear up to now. Particularly, how to keep the aerodynamic performance similarity in the case of geometry dissimilarity has not been explored yet. Here we propose a geometry scaling technique in which 3D parametric geometry modeling and 1D mean-line analysis are combined to define a series of key redesign criteria in order to keep as minor as possible variations in the aerodynamic performance of multi-stage compressor. The proposed technique is applied to redesign a 3.5 stage transonic axial-flow compressor with two geometry scaling strategies, i.e., adjustable blade/vane numbers and inlet guide vane (IGV)-rotor-stator axial spacing. The key aerodynamic parameters determined by the key redesign criteria are then quantified to verify the feasibility and accuracy of the proposed technique by computational fluid dynamics (CFD) simulations. The results show that by keeping the similarity of key redesign criteria, that is, blade solidity, stagger angle, inlet blade angle, outlet blade angle and thickness to chord distribution, the similarity of key aerodynamic parameters (incidence angle, deviation angle, loss coefficient and aerodynamic blockage) and thus of overall aerodynamic performance (mass flow rate, pressure ratio and efficiency) can be generally guaranteed. The variations of aerodynamic performance in the case of geometry dissimilarity with different blade/vane numbers and IGV-rotor-stator axial spacing are generally within 1%. Compared against the similarity of the other key aerodynamic parameters, a certain dissimilarity of aerodynamic blockage is observed, impacting slightly on the mass flow rate and thus stall margin. The present work is of effective assistance to improve the redesign methodology for multi-stage axial-flow compressors.

Stability management of high speed axial flow compressor stage through axial extensions of bend skewed casing treatment

This paper presents the experimental results to understand the performance of moderately loaded high speed single stage transonic axial flow compressor subjected to various configurations of axial extensions of bend skewed casing treatment with moderate porosity. The bend skewed casing treatment of 33% porosity was coupled with rectangular plenum chamber of depth equal to the slots depth. The five axial extensions of 20%, 40%, 60%, 80% and 100% were used for the experimental evaluations of compressor performance. The main objective was to identify the optimum extension of the casing treatment with reference to rotor leading edge which results in maximum stall margin improvements with minimum loss in the stage efficiency. At each axial extension the compressor performance is distinctive. The improvement in the stall margin was very significant at some axial extensions with 4%–5% penalty in the stage efficiency. The compressors stage shows recovery in terms of efficiency at lower axial extensions of 20% and 40% with increase in the peak stage efficiency. Measurements of flow parameters showed the typical behaviors at near stall flow conditions. Hot wire sensor was placed at the rotor upstream in the tip region to capture the oscillations in the inlet axial and tangential velocities at stall conditions. In the absence of casing treatment the compressor exhibit abrupt stall with very high oscillations in the inlet axial and tangential velocity of the flow. The extents of oscillations reduce with bend skewed casing treatment. Few measurements were also performed in the plenum chamber and salient results are presented in this paper.

Improvement of Moderately Loaded Transonic Axial Compressor Performance Using Low Porosity Bend Skewed Casing Treatment

This paper presents experimental results of a single stage transonic axial flow compressor coupled with low porosity bend skewed casing treatment. The casing treatment has a plenum chamber above the bend slots. The depth of the plenum chamber is varied to understand its impact on the performance of compressor stage. The performance of the compressor stage is evaluated for casing treatment and plenum chamber configurations at two axial locations of 20% and 40%. Experimental results reveal that the stall margin of the compressor stage increases with increase in the plenum chamber volume. Hot-wire measurements show significant reduction in the turbulence intensity with increase in the plenum chamber volume compared to that with the solid casing at the stall condition. At higher operating speeds of 80% and at 20% axial coverage, the stall margin of the compressor increases by 20% with half and full plenum depth. The improvement in the peak stage efficiency observed is 4.6% with half plenum configuration and 3.34% with the full plenum configuration. The maximum improvement in the stall margin of 29.16% is obtained at 50% operating speed with full plenum configurations at 40% axial coverage.

Inverse Design of 3D Multistage Transonic Fans at Dual Operating Points

Semi-inverse design is the automatic recambering of an aerofoil during a computational fluid dynamics (CFD) calculation in order to achieve a target lift distribution while maintaining thickness, hence, “semi-inverse.” In this design method, the streamwise distribution of curvature is replaced by a streamwise distribution of lift. The authors have developed an inverse design code based on the method of Hield (2008, “Semi-Inverse Design Applied to an Eight Stage Transonic Axial Flow Compressor,” ASME Paper No. GT2008-50430), which can rapidly design three-dimensional fan blades in a multistage environment. The algorithm uses an inner loop to design to radially varying target lift distributions, an outer loop to achieve radial distributions of stage pressure ratio and exit flow angle, and a choked nozzle to set design mass flow. The code is easily wrapped around any CFD solver. In this paper, we describe a novel algorithm for designing simultaneously for specified performance at full speed and peak efficiency at part speed, without trade-offs between the targets at each of the two operating points. We also introduce a novel adaptive target lift distribution, which automatically develops discontinuous changes of calculated magnitude, based on the passage shock, eliminating erroneous lift demands in the shock vicinity and maintaining a smooth aerofoil.

Experimental Investigation of a Two Stage Transonic Axial-Flow Compressor

Experimental Investigation of a Two Stage Transonic Axial-Flow Compressor