Research using numerical simulation methods on the aerodynamic performance of the high bypass ratio civil engine exhaust system for civil aircraft has been launched in this paper. The influence of flight conditions and aircraft integration on the aerodynamic performance of the exhaust system is systematically analyzed. The result shows that flight Mach number manifests an inhibition influence on the mass flow of the exhaust system and a positive impact on the thrust of the exhaust system. Flight altitude has a negative influence on the mass flow and the thrust of the exhaust system. Aircraft body integration represents a gain on the exhausting flow and an opposite influence on the thrust.

In order to study the effect of inlet paired swirl distortion on the performance and stability of compressor, a full-annulus unsteady numerical simulation of the effects of swirl distortion on performance and stability of transonic rotor Rotor 37 is carried out in this paper. The result shows that: the rotor performance deteriorates and the stable working range decreases under the effect of vortex flow distortion; the circumferential side flow caused by vortex swirl distortion leads to the uneven circumferential distribution of flow parameters at the inlet and the outlet of the compressor; the rotor blade has a very significant attenuation effect on the swirl distortion, and the swirl distortion intensity decreases along the axial direction; the circumferentially inhomogeneous distribution of rotor angle causes excessive local aerodynamic load on the rotor, and high intensity leakage vortices and excitation vortices are generated at the blade tip. The circumferentially unevenly distributed swirl angle causes excessive local aerodynamic load of the rotor, resulting in high-intensity leakage vortex and shock wave interference in the blade tip area, which severely blocks the flow path and eventually leads to unstable flow in the Rotor37.

民用飞机通过机载大气数据传感器测量飞机所处环境下的大气总压与静压,并计算空速.为研究民用飞机总压传感器布局位置的气动特性,采用CFD数值计算手段得到不同迎角和侧滑角工况下机头附近的流场,通过候选总压探头布局位置的局部气流角度和探头总压损失曲线获得不同工况下的总压损失和空速误差.当局部角度曲线斜率低且各曲线聚集时,是总压探头的理想最优布局位置,探头局部气流角对迎角和侧滑角均不敏感,能同时保证横向和纵向气动特性最优;通过探头位置的局部气流角度并结合风洞试验获得的总压损失系数分布曲线进行总压损失预估,该方法对总压损失的预测准确有效,预测精度满足要求,可用来进行空速误差估算;总压探头布局位置的选取需结合民用飞机实际运行的侧滑角/迎角包线综合判断.

Blended-wing-body aircraft is a new type of aircraft form that has been used in military aircraft. The potential for application in civil aircraft is being explored. Due to its special shape, the blended-wing-body aircraft is difficult to adopt the conventional under-wing engine mounting form, and usually applies the back braced mounting. The engine mounted on the back could cause surge and separation problems, which affects the flow field quality of the whole aircraft. In this paper, the effect of height and axial mounting position on the overall flow field of the aircraft is determined by modelling and calculation of the blended-wing-body aircraft. The effect of airflow separation is analysed. The engine axial mounting position is discovered to have a significant effect on the elimination of flow separation.

附面层边缘通常取在速度达到主流速度0.99倍的位置,而复杂流场中主流流动往往并不均匀,给附面层边缘的准确识别造成了困难.为解决此问题,提出了用“参考主流”代替实际主流识别附面层边缘的方法:通过零剪切力滑移壁面边界条件下数值模拟得到不受附面层干扰的参考主流,在根据附面层定义确定附面层边缘时以该参考主流中的速度代替实际的主流速度.通过斜楔压缩和弯曲压缩两个超声速压缩流场对该识别方法进行了验证,所得到的斜楔压缩出口截面上附面层厚度与采用实际主流速度判断得到的厚度相对误差仅4.1％.根据该方法的识别结果对弯曲压缩型面设计进行附面层修正后,弯曲激波高度与无黏设计值之间的误差从修正前的2.0％降低至0.3％,压缩面末端压力的相对误差从修正前的6.6％降低至2.3％.该方法避免了指定主流速度的主观性,识别结果较为准确.