Research Of End-flow Effect On Compressor Stability With Inlet Total Pressure Distortions

Aircraft engine is a strategic product which plays a significant role in defense industry and domestic economic, and its related technology is the core of21st century propulsion system, directly influences the defense, energy, secure and industry of a country. Therefore, advanced aircraft engine technology has become an essential mark for measuring the level of national science and technology, industrial advancement, military force and even the comprehensive national power, and meanwhile the priority in the development of technology and military’industry for each country. With the rise of speed, altitude, increasing maneuverability and the use of missile weapon of military aircraft, the influence of distortion is more and more highlighting, and the compatibility of inlet and engine when exposed in practice is more and more serious. The limitation of early stage design which based on the uniform inlet condition is becoming larger, no longer suitable to the high performance engine design. Especially the4S requirements of the fourth fighter, which is "Super maneuverability","Super Sonic Cruise","Stealth" and "Superior Avionics for Battle Awareness and Effectiveness", make the compressor be under distorted inlet conditions for most cases, putting forward more demands on its stability and anti-distortion ability. It is can be foreseen that distortion has become one of the key research areas. Understanding of the unsteady flow mechanism and flow characteristic changes with distorted inflow in compressor is significant and realistic to the design system of compressor.Sponsored by the NSF-Young Fund (51006014), a single stage transonic axial compressor is studied under distorted inlet conditions in this thesis. In order to gain a more accurate distorted inlet, numerical method is used to investigate the distorted flow field generated by a baffle distortion generator, and obtains the inlet boundary condition of the compressor. The result shows the distortion degree of compressor inlet has an approximately linear relation to inflow total pressure. The total pressure distribution behind the baffle could be simplified into high and low pressure regions (each uniformly distributed). And the total pressure distribution of compressor inlet section differs greatly with that after the baffle:the low pressure region expanses as depth of baffle increases thus could not be simplified to high and low pressure regions. Meanwhile, due to the complicated unsteady three dimensional viscous flows inside the compressor, even under the uniform inflow condition, the flow in the passages would involve various phenomena as it develops, such as boundary layer transition, separation and reattachment. In order to accurately analyze the influence caused by distortion, this dissertation carries out full annuls unsteady three-dimension numerical simulation under distorted inlet conditions as well as the uniform inlet condition. Result show that under the uniform inflow condition, the flow in the compressor passage distributes uniformly, and shows a good periodicity. At the maximum efficiency point, a detached shock wave forms at the rotor leading edge and a normal shock wave generates in the passage. A small separation forms at hub and shroud corner near trailing edge of stator suction. Induced by the rotor wake, time-varying vortex forms at the stator hub near leading edge. Near stall point, only one shock wave exists in rotor passage and a wide range of circumferential flow separates near the hub in stator. Shock wave at the rotor tip and separation near the stator hub are important flow characteristics in the compressor, especially at the near stall point. Therefore, this dissertation focuses on studying the changes of shock wave in rotor and separation in stator to analyze flow in the compressor under distortion conditions.Before analyzing the influence of distortion to the flow field of compressor, the transfer law of distortion inside the compressor is studied in this dissertation firstly: after the rotor, the location of the lowest total pressure mitigates in the opposite rotating direction while the location of maximum losses mitigates in the rotating direction. Through comparably studying the numerical result of the uniform inflow condition and distorted inflow condition at various degrees, this dissertation studies the influence of compressor performance and internal flow field at different distortion degree. The result shows the compressor performance deteriorates obviously and the stability margin declines when the distortion degree increases. Because of the distortion, there exist two pressure gradients in the flow:axial pressure gradient and circumferential pressure gradient. The axial pressure gradient mainly leads to the mitigation of shock wave in rotor passage while the circumferential pressure gradient mainly causes the change of the flow angle by generating the circumferential velocity. Additionally, this dissertation analyzes the change rule of shock wave in the rotor passage and separation in the stator passage under distorted condition in detail. At last, compared to the flow field of compressor at near stall point in uniform condition, this dissertation analyzes the main source of the compressor stall.Based on above results, the internal flow of compressor at different specific speed under distorted inlet is also studied in this dissertation. Firstly, according to the distribution of inlet total pressure, the compressor studied in this thesis can be divided into two sub-compressors with different inlet total pressure. Aerodynamic parameters of the compressor at outlet are close to that of the sub-compressor which has a larger proportion, while the other sub-compressor can be seen as an external stimulus, namely "distortion zone". With the increase of specific speed, the effect of distorted inflow to the aerodynamic performance of the compressor becomes more and more significant. When the distortion intensity is low, the compressor performance deteriorates obviously at high specific speed; when the distortion intensity is high, the compressor performance deteriorates more significantly at low specific speed. Additionally, comprehensive analysis of the shock wave in rotor and flow capacity of stator under different speed conditions shows that at low specific speed, the shock wave is far from the leading edge, easier to put away from the rotor passage, and thus unable to form a stable shock wave structure, causing the compressor instability. At the same time, the stator has a smaller separation and a better flow capacity. Therefore, under low specific speed, the instability of compressor is mainly caused by the break of shock wave structure. While at high specific speed, the separation range and strength in stator passage significantly increase, thus the flow capacity dramatically weakens, and the passage is blocked seriously. Yet this time the distance between shock wave and leading edge is close, and the position of shock wave hitting the adjacent rotor blade suction is far from the leading edge, thus the shock wave is more stable. Therefore the instability of compressor is caused by passage blockage in stator.