Experimental And Numerical Investigation On Endwall Flow Of Cantilevered Stator In Axial Compressor

In axial compressor design process, there are two choices for stator configuration: cantilevered versus shrouded. By using different stator configurations, the endwall flow will be changed, thus the compressor aerodynamic performance will be influenced. However, there is no consensus about the influence mechanism. The associated endwall flow structure change mechanism is still needed to be further investigated to exploit the potential of the compressor performance improvement. In this paper, a linear compressor cascade is used to experimentally and numerically investigate the flow structure change mechanism induced by choosing different stator structure in a large operation range. The shrouded stator configuration is simulated by the non-clearance cascade, while the cantilevered stator configuration is simulated by clearance cascades. The influence of different endwall flow structures on cascade aerodynamic performance is compared to give a theoretical support for optimization of endwall flow field and stage matching design.In the first part of this paper, the non-clearance cascade aerodynamic performance and the cascade internal flow structure are experimentally investigated in a large range of incidences. At highly negative incidence angle condition, a pair of three-dimensional separation vortexes near the pressure-side leading edge is originated from the end wall, generating a high-loss area at cascade outlet midspan. At design condition, a slight corner separation exists at the endwall, and the cascade loss almost remains unchanged as the incidence angle increases. At corner stall condition, the endwall aerodynamic performance is severely deteriorated by the corner stall structure, and the cascade loss increases as the incidence angle increases. The corner stall structure is formed by the flow close to the endwall driven by secondary flow effect and the backflow driven by the trailing edge pressure difference. The increase of the inlet boundary layer thickness and inflow mainstream velocity will make corner stall happen at smaller incidence angle.In the second part, the influence of clearance variation on endwall flow and cascade aerodynamic performance is investigated at three typical operation conditions. At highly negative incidence angle condition, the leading edge separation vortexes exist in different clearance cascades. The cascade total pressure loss decreases with the increase of clearance size, and the clearance side turning ability is slightly increased. At design condition, by bringing in a small clearance (0.2%span), corner separation is first aggravated. As the clearance size continues to increase, the corner separation can be eliminated, and leakage vortex is formed. With the increase of clearance size, the cascade total pressure loss first increases, then decreases and finally goes up again. At corner stall condition, the introduction and increase of clearance can impair and eliminate the corner stall structure, thus the cascade loss decreases and reaches the minimum value with a 0.5%span clearance size. And the clearance side turning ability is also improved. Flow structures at both enwall can influence each other as clearance varies. By oil-flow visualization experiments and numerical simulation, it is proven that the leakage flow at different axial position is originated from areas near the pressure-side leading edge. A low static pressure area exists in the clearance and near the suction-side surface, and it will be shrunk to leading edge as clearance decreases, which reduces the leakage flow after the middle chordwise position in favor of the formation of corner separation. When the corner separation attaches to the blade suction-side surface, the endwall is dominated by corner separation. Otherwise, the endwall flow field is dominated by leakage vortex. The inlet boundary layer thickness variation has little influence on leakage vortex. When the corner separation and leakage flow both exist, the increase of inlet boundary layer thickness will aggravate the corner separation, which impairs the cascade aerodynamic performance. The influence mechanism of clearance variation on enwall flow structure nearly remains the same when the inlet mainstream velocity is lower than 0.45Ma.In the third part, the influence of chordwise stepped non-uniform clearance on the endwall flow at clearance side of a cantilevered stator is numerically studied. It is shown that an optimal stepped non-uniform clearance structure exists to improve the cascade performance. The relative large clearance placed at rear chordwise part shows a better effect than placed at the front part. By experimental and numerical research, it is proven that a proper cascade chamfer can impair corner stall structure at the non-clearance side, which further enhances the cantilevered stator aerodynamic performance.In the fourth part, different endwall flow characteristics are investigated by a hot-wire anemometer at different outlet axial positions. The results show that the outlet high turbulent intensity areas are similar with the high total pressure loss areas. The corner stall structure has a low characteristic frequency, while the leakage vortex and corner separation shows no such characteristic frequency in linear cascade environment. The spectral characteristic of sample data measured at different spatial points in the corner stall structure are quite different from each other, which shows that corner stall structure is anisotropic. While the spectral characteristic of sample data measured at different spatial points in a leakage vortex is similar. When the large corner separation and leakage flow both exist at the small clearance endall area, the unsteady characteristic is dominated by corner separation. The spectral characteristic of corner stall structure attenuates more quickly than the leakage vortex, as the flow structures develop downstream the cascade.