Structure of Tip Leakage Flow in a Forward-Swept Axial-Flow Fan

The rotor-tip flow fields in two rotors of a low-speed axial-flow fan were experimentally and numerically investigated to clarify the mechanism behind modal stall inception. A NACA 65 wing section and a controlled diffusion airfoil were applied to the two rotors. At the small stagger-angle setting for both rotors, which is ten degrees smaller than the design value, the modal disturbance is observed near the peak pressure-rise point, and the rotor blades at the tip stall before the modal disturbance is observed. In the modal stall inception, the interface between the incoming flow and the reversed tip-leakage flow does not become parallel to the leading edge plane, although backflow from the trailing edge initiates near the stall condition. The reversed tip-leakage flow does not spill from the leading edge at the stall condition. Moreover, the tip-leakage vortex breakdown does not occur near or at the stall condition. A three-dimensional separation vortex is induced by secondary flow on the suction surface near the stall condition and develops at the stall condition. It is concluded from these results that the rotor-tip flow fields in the modal stall inception differ from those in the spike stall inception and that the three-dimensional separation vortex induced by the secondary flow influences the initiation of modal disturbance.

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Effects of Tip Leakage Flow on the Aerodynamic Performance and Stability of an Axial-Flow Transonic Compressor Stage

Energies ◽

10.3390/en14144168 ◽

2021 ◽

Vol 14 (14) ◽

pp. 4168

Author(s):

Botao Zhang ◽

Xiaochen Mao ◽

Xiaoxiong Wu ◽

Bo Liu

Keyword(s):

Boundary Layer ◽

Shock Wave ◽

Axial Flow ◽

Leading Edge ◽

Rotating Stall ◽

Tip Clearance ◽

Leakage Flow ◽

Tip Leakage Flow ◽

Tip Leakage ◽

Transonic Compressor

To explain the effect of tip leakage flow on the performance of an axial-flow transonic compressor, the compressors with different rotor tip clearances were studied numerically. The results show that as the rotor tip clearance increases, the leakage flow intensity is increased, the shock wave position is moved backward, and the interaction between the tip leakage vortex and shock wave is intensified, while that between the boundary layer and shock wave is weakened. Most of all, the stall mechanisms of the compressors with varying rotor tip clearances are different. The clearance leakage flow is the main cause of the rotating stall under large rotor tip clearance. However, the stall form for the compressor with half of the designed tip clearance is caused by the joint action of the rotor tip stall caused by the leakage flow spillage at the blade leading edge and the whole blade span stall caused by the separation of the boundary layer of the rotor and the stator passage. Within the investigated varied range, when the rotor tip clearance size is half of the design, the compressor performance is improved best, and the peak efficiency and stall margin are increased by 0.2% and 3.5%, respectively.

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A Proposal for Passive Wall Treatment Applied in High Performance Axial Flow Compressors

Volume 2A: Turbomachinery ◽

10.1115/gt2020-15238 ◽

2020 ◽

Author(s):

Rubén Bruno Díaz ◽

Jesuino Takachi Tomita ◽

Cleverson Bringhenti ◽

Francisco Carlos Elizio de Paula ◽

Luiz Henrique Lindquist Whitacker

Keyword(s):

Stokes Equations ◽

Axial Compressor ◽

Axial Flow ◽

Tip Clearance ◽

Leakage Flow ◽

Smooth Wall ◽

Tip Leakage Flow ◽

Viscous Effects ◽

Stall Margin ◽

Tip Leakage

Abstract Numerical simulations were carried out with the purpose of investigating the effect of applying circumferential grooves at axial compressor casing passive wall treatment to enhance the stall margin and change the tip leakage flow. The tip leakage flow is pointed out as one of the main contributors to stall inception in axial compressors. Hence, it is of major importance to treat appropriately the flow in this region. Circumferential grooves have shown a good performance in enhancing the stall margin in previous researches by changing the flow path in the tip clearance region. In this work, a passive wall treatment with four circumferential grooves was applied in the transonic axial compressor NASA Rotor 37. Its effect on the axial compressor performance and the flow in the tip clearance region was analyzed and set against the results attained for the smooth wall case. A 2.63% increase in the operational range of the axial compressor running at 100%N, was achieved, when compared with the original smooth wall casing configuration. The grooves installed at compressor casing, causes an increase in the flow entropy generation due to the high viscous effects in this gap region, between the rotor tip surface and casing with grooves. These viscous effects cause a drop in the turbomachine efficiency. For the grooves configurations used in this work, an efficiency drop of 0.7% was observed, compared with the original smooth wall. All the simulations were performed based on 3D turbulent flow calculations using Reynolds Averaged Navier-Stokes equations, and the flow eddy viscosity was determined using the two-equation SST turbulence model. The details of the grooves geometrical dimensions and its implementation are described in the paper.

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