Abstract
Multistage compressor is the most important constituent of gas turbines used in land, naval and aeronautical applications. Overall performance of such machinery depends mainly on the axial compressor performance. Due to the relative motion between rotor and stator blades, the flow field in this machinery is highly unsteady. Furthermore several technological effects like tip clearances, complexity of the blade shapes, variation of axial distance between stator and rotor, seal leakages and cooling holes among others complicate the machine. Therefore the study of a complicated, strongly three-dimensional flow field inside a compressor is considered to be one of the most difficult tasks to be performed by a CFD expert. The present work is the extensive numerical study of the effect of: (1) tip clearance of rotor blades and (2) the axial gap between rotor and stator on the overall performance of a multistage axial compressor. A commercial software package is used for this study. Reynolds-averaged Navier–Stokes equations are solved using Spalart–Allmaras model. A number of steady-state viscous flow simulations were run for both the tip clearance effect and different axial gaps between stator and rotor. All simulations were performed for the first stage, i.e. Stator–Rotor–Stator. Simulations were carried out with coarse, medium and fine meshes to find an optimum, mesh-independent solution. It has been found that larger tip clearance has a detrimental effect on the stage pressure ratio and efficiency of a multistage axial compressor. Similarly there exists a certain distance ratio between the stator1–rotor and rotor–stator2, where stage performance is optimum. Overall performance characteristics obtained through simulation for both the tip clearance and axial gap variation were also compared with the experimental studies and found to be in good agreement.
Highlights
Apart from various industrial applications, the multistage axial compressor is the most important constituent of all gas turbine engines for land and aeronautical applications
Measured data of this study showed that the axial spacing between an upstream stator and downstream transonic rotor has a noteworthy effect on stage performance
Outlet flow field of the multistage axial compressor with bowed stator stages was measured by Huawei et al [20] using five hole probe under near stall conditions when rotor–stator axial gap was reduced by 33%, 67% and 100% of the original gap
Summary
Apart from various industrial applications, the multistage axial compressor is the most important constituent of all gas turbine engines for land and aeronautical applications. Yoon et al [17] studied the effect of stator hub configuration on the aerodynamic performance They investigated the leakage loss and concluded that the design of clearance/span of the axial compressor is vital in optimizing the performance. Outlet flow field of the multistage axial compressor with bowed stator stages was measured by Huawei et al [20] using five hole probe under near stall conditions when rotor–stator axial gap was reduced by 33%, 67% and 100% of the original gap They found that the overall performance increases within the main flow region as the rotor–stator axial gap of the compressor is decreased. The present work is the extensive numerical study of the effect of (1) the tip clearance of rotor blades and (2) the axial gap between rotor and stator on the efficiency of a multistage axial compressor. The simulations were run for many coarse and fine meshes in order to find an optimum, meshindependent solution
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