Flow Control Effect of Spanwise Distributed Pulsed Arc Discharge Plasma Actuation on Supersonic Compressor Cascade Flow
- PDF / 2,349,673 Bytes
- 11 Pages / 595.22 x 842 pts (A4) Page_size
- 65 Downloads / 224 Views
https://doi.org/10.1007/s11630-020-1382-3
Article ID: 1003-2169(2020)00-0000-00
Flow Control Effect of Spanwise Distributed Pulsed Arc Discharge Plasma Actuation on Supersonic Compressor Cascade Flow SHENG Jiaming1, WU Yun1,2, ZHANG Haideng1,2*, WANG Yizhou1, TANG Mengxiao1 1 Science and Technology on Plasma Dynamic Laboratory, Air Force Engineering University, Xi’an 710038, China 2 Institute of Aero-engine, School of Mechanical Engineering, Xi’an Jiaotong University, Xi’an 710049, China © Science Press, Institute of Engineering Thermophysics, CAS and Springer-Verlag GmbH Germany, part of Springer Nature 2020
Abstract: To achieve efficient control of supersonic compressor cascade flow, a type of spanwise distributed pulsed arc discharge plasma actuation (PADPA) was designed. To simulate the influences of PADPA on the flow field, a phenomenological model was established. Then, the flow control effects of PADPA on supersonic compressor cascade flow were researched numerically. The results show that under low static pressure ratio condition, the compressive wave induced by PADPA reduced the intensity of the passage shock wave, which eventually reduced shock wave loss. It was also found that PADPA produced an adverse pressure gradient (pre-compression effect) around the actuation location, which reduced the strength of the high adverse pressure gradient induced by the passage shock wave. The airflow on both sides of the actuation location was accelerated by PADPA owing to the spanwise distributed layout. Thus, it improved the ability of the boundary layer to resist the effect of the adverse pressure gradient and reduced the separation zone. Consequently, the total pressure loss was reduced by 6.8%. Under high pressure ratio condition, the effect of PADPA on the suction side controlling the large separation of the boundary layer was insignificant. The total pressure loss also increased slightly.
Keywords: plasma, flow control, supersonic cascade, shock wave/boundary layer interaction, numerical simulation
1. Introduction With the continuous development of technology, the aircraft engine rotating speed and the flow velocity relative to the blades continue to increase, even reaching transonic or supersonic speed. Under supersonic inlet conditions, the static pressure increases owing to the shock wave structure arising in the blade passage. However, certain new problems have been encountered while applying this method. First, as the rotational speed of the compressor and intensity of the shock wave continue to increase, shock wave loss becomes larger. Second, because of the interaction between the shock Received: Jun 09, 2020
AE: KAN Xiaoxu
wave and boundary layer, the boundary layer separates under a strong adverse pressure gradient, resulting in an increase in viscosity loss [1]. Third, the shock wave/ boundary layer interaction induces low-frequency shock wave oscillation. Then, the low-frequency oscillation of aerodynamic load will lead to vibration of the mechanical structure which might endanger flight safety [2
Data Loading...
