【作者】 高海洋；

【导师】 钟兢军； 孙鹏；

【作者基本信息】 大连海事大学 ， 轮机工程， 2014， 博士

【摘要】 航空发动机是国防工业和国民经济具有重大作用的战略产品,其相关技术是21世纪动力系统的核心技术,直接影响着一个国家的国防、能源、安全和工业竞争能力。因此,先进的航空发动机技术已经成为一个国家科技水平、工业先进水平、军事实力、乃至综合国力的一个重要标志,同时也是各国科技及军事工业优先发展的领域和重点研发的对象。随着军用飞机飞行速度、高度的不断提高,机动性的不断增加以及导弹武器的使用,畸变的影响变得越来越突出,实践中暴露出来的进气道/发动机相容性问题也越来越严重。早期基于均匀进口条件进行的设计所暴露的局限性越来越大,己不再适合高性能发动机的设计。特别是第四代战机的4S要求,即“超机动性”、“超音速巡航”、“隐身能力”和“超高效空战航电设备”,使得压气机大部分时间都在进口畸变条件下工作,对压气机的稳定性和抗畸变能力提出了更高的要求。可见,畸变问题已经成为航空发动机领域重点的研究方向之一,认识畸变条件下压气机内部非定常流动机理以及掌握流场结构变化对畸变来流的响应机制,进而指导压气机的设计具有迫切的现实意义。本文在国家自然科学基金青年基金项目(51006014)资助下,以单级跨声速轴流压气机为对象,研究总压畸变条件下的内部流动特性。为了更加准确的给定畸变进口,本文首先采用数值方法对插板式畸变发生器产生的畸变流场进行了研究,并获得压气机进口边界条件。结果表明：压气机进口截面畸变度与来流总压近似成线性关系。插板后的总压分布可以简化为两个各自均匀的高/低压区域。而压气机进口截面的总压分布与插板后相比存在较大差异,低压区范围随插板深度增加而增大,不能简单简化为两个各自均匀的高/低压区域。同时,由于压气机内部本身就是十分复杂的三维非定常黏性流动,即使在均匀来流条件下,随着气流在流道内的发展,叶栅流道内也会涉及到各种各样的流动问题,如附面层的转捩、分离和再附等。为了更加准确分析由畸变带来的影响,本文在进行畸变条件下全周非定常三维数值模拟工作的同时,还进行了均匀进口来流条件下压气机全周非定常三维数值模拟工作。结果表明：均匀进口条件下压气机内各流道流场分布均匀,呈现出良好的周期性。最高效率点,动叶前缘形成一道脱体激波,通道内形成一道正激波。静叶吸力面尾缘叶顶、叶根位置形成较小的分离。静叶根部吸力面前缘附近,在动叶尾迹作用下分离形成一个旋涡结构,且随时间不断变化。近失速点,动叶流道内仅有一道激波,静叶各流道叶根附近表现为大范围分离流动。压气机动叶叶顶激波和静叶叶根分离是压气机内,尤其是近失速状态下,重要的流动特征。因此,本文将主要针对动叶内激波和静叶中分离的变化来进行畸变条件下压气机内部流场的分析研究。然后,本文先分析了畸变在压气机中的传递规律。研究发现,畸变经过动叶后,总压最低值位置逆旋转方向偏转；损失最大值位置顺旋转方向偏转。通过对比均匀进口和不同畸变度条件下的数值结果,研究了畸变度对压气机性能及内部流场的影响。结果表明：随畸变度提高,压气机性能下降明显,稳定工作流量范围变窄。同时由于畸变的存在,流场内产生了两个方向的压力梯度,即轴向压力梯度和周向压力梯度。轴向压力梯度主要导致动叶流道内激波位置发生偏移,而周向压力梯度产生的周向分速度主要引起气流角发生变化。在此基础上,详细分析了不同状态点时动叶流道内激波和静叶流道内分离的演变规律,阐述了引起动静叶流场参数变化的机理。并通过与均匀进口条件压气机近失速点流动状态对比,分析了畸变条件下压气机失稳诱因的主要来源。最后,本文在上述研究的基础上,研究了不同比转速时畸变对压气机流场的影响。根据压气机进口总压的分布情况,将本文所研究的压气机分成两个不同来流总压的子压气机。压气机出口气动参数趋近于所占比例较大的子压气机出口的气动参数,而另一子压气机则可以看作外部激励,即“畸变区”。随比转速增大,激励次数增加,畸变来流对压气机气动性能的影响越来越明显。研究表明,低畸变度、高比转速条件下,压气机性能下降更明显。高畸变度、低比转速条件下,压气机性能下降更明显。综合分析不同比转速条件时,动叶激波位置和静叶通流能力发现：低比转速时,动叶内激波距前缘较远,激波更容易被推出动叶流道,无法形成稳定的激波结构,造成压气机失稳。而此时静叶内流动分离较小,通流能力较好。因此,低比转速条件下,压气机的失稳诱因主要是动叶内激波结构的破坏。而高比转速时,静叶流道内流动分离范围、强度均明显增大,静叶的通流能力减弱,静叶流道内堵塞严重。而对应此时动叶内激波距离前缘较近,激波打到相邻叶片吸力面的位置也更远离前缘,激波结构的稳定性相对较好,即此时压气机的失稳主要源于静叶流道内流动的堵塞。更多还原

【Abstract】 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.更多还原