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.

为了研究民用飞机进气道在起飞低速大迎角状态下的流场特征和性能,对设计马赫数为0.785的进气道进行了风洞实验和数值计算,来流马赫数为0.2,迎角变化为0°～25°,流量系数为0.29～2.07.研究结果表明:在起飞工况条件下,进气道正常工作迎角可达到25°;在起飞单发失效工况条件下,进气道外罩上流动分离迎角在13°～16°;在大迎角工作条件下内流未产生分离时,出口气流畸变受发动机流量变化影响较小.通过研究结果分析了进气道在低速大迎角状态下的流场特征及其随迎角、流量的变化规律.数值计算和风洞实验结果的对比表明,在没有流动分离时两者基本一致;但是对于外流分离迎角的预测,两者有3.差异,表明风洞实验依然是考核进气道性能的必要手段.

For the situation that the pitching moment of a civil fuselage-mounted aircraft shows to pitch up,the paper studied a new type of improvement measures—to truncate the inner slat.The numerical simulation method was used to study the aerodynamic characteristics of the basic landing configuration and the improvement configuration.The flow characteristics of the two configurations under the high angle of attack are given,and the flow mechanism of the improvement measures is analyzed.The influence of the improvement measures on the intake of the engine is also evaluated.The results show that: the low-speed pitching moment shape depends on the separation position of the wing;separation from the inside wing makes the aircraft"rise";separation from the outer wing makes the aircraft "head",and separation from the inside wing makes the aircraft "bow";the pitching moment provided by the horizontal tail determines the total pitching moment;the main reason for the aircraft "bow"is that the separation from the inside wing reduces the downwash of the horizontal tail from the wing; the improvement measures do not affect the engine intake of the fuselage-mounted aircraft when the angle of attack is not greater than the stall angle of attack.

为准确获取大涵道比发动机气动特性,根据民用飞机研制需求,对民用干线飞机的发动机短舱三维流场进行数值模拟,对比了通流短舱和真实发动机短舱流场特征,分析了通流短舱的气动特性和阻力修正方法.结果表明:大涵道比发动机通流短舱的流场特征、气动力和真实发动机较为接近,但是由于发动机内流通道气动力一般计入推力和无法完全模拟发动机进排气效应,有必要对通流短舱进行考虑内部阻力在内的气动特性修正;通过对大涵道比发动机推阻力划分理论进行分析,获得了与真实发动机一致的通流短舱气动力分解方法,并可以较好地应用到通流短舱内部阻力修正上;通过对通流短舱阻力修正(含内阻修正和溢流阻力修正),在巡航状态下修正后的通流短舱与发动机动力短舱阻力特性差异降低了约46％;在高速巡航状态,阻力修正后的通流短舱气动特性可以更好地反映真实发动机工作的影响,但依然有必要对发动机动力喷流影响开展进一步的分析工作.

In order to improve the stability of the implicit higher order discontinuous Galerkin (DG)method,an implicit GMRES method based on analytic exact Jacobian matrix is developed for simulating compressible laminar and turbulent flow.In the process of the GMRES,Jacobians of inviscid and viscous flux,which are solved by the chain rule,are used for the LU-SGS preconditioning and GMRES matrix-vector product. The source of the Spalart-Allmaras turbulence model is modified to avoid unphysical solution that occurs at implicit derivative.The developed method is validated by benchmark laminar and turbulent flow cases.The results show that the implicit GMRES method based on exact Jacobian possesses significant advantages in reducing computational cost and improving the stability of the implicit high order discontinuous Galerkin method.

为准确分析并确定飞机气动力从而获取飞机气动特性,根据民用飞机研制和性能飞行试验的研究需求,采用数值模拟方法对飞机和发动机带动力三维流场进行了计算,分析了安装和非安装状态下发动机附近流场和其推力参数的变化,初步获得了发动机安装效应对尾吊式民机推力预测的影响.结果表明:本文采用的基于流管假设的推阻力划分方法和数值模拟分析方法,可以获得发动机安装前后的总推力、净推力、安装推力和各推力分量,其结果与发动机性能模型预测基本一致;在带动力条件下,对飞机可用推力的预测需仔细分析安装效应对发动机安装推力的影响,和非安装状态不同,安装状态下喷管气流易受机体/机翼/吊挂流场干扰,其上产生较为明显的压缩-膨胀-再压缩过程;对发动机安装和非安装状态内外涵喷管流动分析表明,出口气流的压力损失和摩阻差异可能是导致推力分量产生变化的主要原因.

In order to study the propulsion effects on the aerodynamic performance of a civil aircraft by us?ing a Turbine Powered Simulator (TPS) in the wind tunnel experiment,numerical simulations were performed and the aerodynamic characteristics were investigated on the TPS nacelle as well as on the nacelle of a real large bypass ratio engine at low speed conditions of takeoff and approach. The lip and cowl geometry of the TPS nacelle was kept the same as that of real engine. The results suggest that some flow-field features of the TPS nacelle are different from the field of engine nacelle in terms of flow capture characteristics,resulting in a slightly different drag coefficient. When the inlet is at subcritical condition,the drag coefficient of TPS is 1.7 counts higher than that of the real engine. In addition,the flow on the upper cowl of TPS nacelle would be easier to separate due to the lower mass flow ratio at high angle-of-attack (AOA) than the real nacelle, therefore, the stall AOA for TPS is about 1.0° smaller. When the inlet is at supercritical condition,due to the variation of stagnation locations and contractions of flow tube with AOA, the drag coefficient of TPS nacelle is higher than that of real engine at AOA about 0°~20° , while the state is reverse at AOA about 20°~30° . Although the differences of drag coeffi?cients are up to 1.8 counts, the aerodynamics characteristics in general are similar between the two. It can be concluded that even TPS nacelle with the same lip and cowl geometry could capture most of the flow features around a real engine. Certain corrections to the drag coefficient of the TPS would nonetheless make the wind tun?nel results more accurate.

根据民用飞机动力装置/机体适航验证研究需要,结合三维有限元碰撞冲击仿真,进行了外物损伤(FOD)条件下发动机短舱内/外流数值模拟分析,初步获得了外罩变形对短舱气动特性的影响.结果表明:结合三维有限元碰撞冲击仿真进行的民用飞机短舱气动特性数值模拟,可以较好地分析FOD对发动机短舱内/外流气动特性和安全性能的影响;在内流方面,外罩和唇口的变形导致了巡航状态和低速大流量等状态短舱内流品质降低,一定程度上影响了进气道流通能力并进一步降低了发动机效率;在外流方面,外物损伤变形导致的气流分离致使局部唇口的前缘吸力丧失,这使得FOD短舱的阻力系数始终比光滑短舱的大,但是当外罩流动均处于失速状态时,两者的阻力特性差异降低.通过对严重影响短舱气动特性或飞机安全性的FOD进行评估,表明该研究成果可以为短舱结构设计的优化和民用飞机安全性分析提供技术依据与建议.

结合民用飞机动力装置/机体集成研究的具体需要,采用数值模拟方法对大涵道比涡扇发动机进气道的侧风畸变特性进行研究,给出了进气道在侧风条件下工作的流场特征,分析了导致侧风畸变的流动机理.研究结果表明:进气道的侧风畸变主要受到典型的气流分离与再附以及地面吸入涡等复杂流动现象的影响;当进气道内部气流出现分离时,侧风畸变会随侧风速度的增加而突然加剧;进气道在近地面工作状态,产生的地面吸入涡现象可能会使短舱的侧风容限明显降低.此外,侧风状态的进气畸变特性随着发动机流量的变化,进气道内部气流的分离与再附存在差别,这导致了进气畸变指数呈现迟滞变化.

Inlet distortion can cause a decrease in the performance of the commercial transport aircraft engines due to the stall and instability which would have significant impact to the economy and safety of the aircraft. Inlet distortion under crosswind for a high bypass ratio turbofan engine is numerically simulated motivated by the need of engine/airframe integration of a commercial transport aircraft. The inlet flow field, characteristics and underlying flow physics are investigated. The investigation suggests that: A sudden increase in the distortion index occurs as the onset of the internal flow separation; Inlet distortion will deteriorate with increasing crosswind speed. Flow separations, reattachments and ground vortex are found to be the root causes for the crosswind distortion. When ground effect is taken into account, it will lead to inlet distortion and its effect will intensified with higher crosswind. The tolerance of crosswind would be reduced as the flowfield affected by the ground vortex phenomenon. The results and methods provide credibility which will technically benefit the engine/airframe integration, experiments as well as corresponding flight tests.

Hypersonic vehicle goes through different kinds of flight condition, experiencing several disturbances. If control system fails to adjust the pose of vehicle, the pitching attitude could be transformed or forced to transient oscillation. The hypersonic inlet performances with steady angle-of-attack are well understood, however, the characteristics with dynamic angleof-attack have received few attentions. In this study, wind tunnel experiment was conducted on a side wall compression scramjet inlet at Mach 3.85, with the range of angle-of-attack 0.08.2 degree and the frequency up to 10.4 Hz. The present work suggests that: In the quasisteady experiment, the inlet would work in unstarting or restarting mode at critical values, then the flow field feature of inlet was significantly altered which was also analyzed by numerical simulation. In the dynamic process, the inlet goes through a unstart-restart process and a hysteresis effect was found which was more remarkable in time of inlet was unstarting; The value of unstarting angle-of-attack would increase and the restarting value would decrease with the rising of frequency in the oscillation.

In order to research the influence of dynamic angle-of-attack on the performance of hypersonic inlet, wind tunnel experiments were conducted on a sidewall compression scramjet inlet with Ma = 3.85, range of angle-of-attack 0°-8.2° and frequency up to 10.4 Hz. The present work suggests that： with the high angle of-attack, the inlet could work in unstart mode, and flow field features are significantly altered and also analyzed by a computational fluid dynamics （CFD） solver. In the dynamic process of angle-of-attack, the inlet could go through a start unstart-restart process hut the pressure on the wall has the hysteresis effect, especially when the inlet is unstart. With the rise of frequency, the value of unstarting angle-of-attack could increase and restarting value could decrease in the oscillation.

For hypersonic vehicle pitching attitude could be altered notably or forced to oscillating by various perturbations.Unsteady numerical simulations were performed on 2-D hypersonic inlet at Ma 4.03.Effects of dynamic angle-of-attack on start/unstart characteristic were analyzed.The investigation suggests that 2D hypersonic inlet will unstart/restart when the pitching attitude alters notably,Then the characteristic and performance of inlet can be significantly changed.For different reduced frequencies,characteristics of flow field and performances of inlet are different.Unsteady flow field feature and the performance lagbehind.The value of unstart angle-of-attack increases and restart value decreases with the rising of reduced frequencies.It is predicted that when the reduced frequency reaches a critical value,depending on the initial angle-of-attack,the inlet will keep start/unstart in whole circle.

In order to study the characteristics of hypersonic inlet response to rapid periodic dynamic angle of attack, numerical simulations are performed on 2D hypersonic inlets with dynamic angles of attack. The effect of dynamic angle of attack on unsteady aerodynamic characteristics is analyzed at a Mach number of 4.03. Two cases of the angle of attack motion are considered. In the first case, the inlet moves upward at a constant angular velocity, while in the second case it repeats a cyclic motion. The investigation suggests that: For dynamic increase of the angle of attack, similar to the steady cases, the characteristics of the flow field and the performance of the inlet are significantly altered, while the unsteady flow field feature and performance lag behind the steady cases. In the first case, the value of the angle of attack at unstart tends to increase with the rising of the angular velocity. In the low flow-speed area near the bottom wall, the impact of the unsteady flow on the aerodynamic characteristics of the inlet is significant: the appearance, development and disappearance of the low flow-speed area caused by the movement of the solid wall would alter the aerodynamic characteristics of the 2D hypersonic inlet.

In the flight, hypersonic vehicle goes through different kinds of flight status, experiencing various of disturbances. If control system fails to adjust the pose of vehicle, the pitching, yawing and rolling attitude could be transformed or forced to oscillate transiently, then the dynamic of angle-of-attack can not be avoided. It’s a question whether hypersonic inlet would keep start and how the performances change. In this paper, unsteady CFD simulation is performed on 2-D hypersonic inlet at Mach 4.03. Effect of dynamic angle-ofattack on starting characteristic is analyzed. The investigation suggests that: The appearance, development and disappearance of low mach number region near solid wall is affected notablely, which would alter the aerodynamic characteristic of inlet. For the dynamic increasing of angle-of-attack, characteristics of flow field and performances of inlet are significantly altered, unsteady flow field features and the performances are lag behind of the steady case. For the oscillating motion, the value of unstarting angle-of-attack increase and restarting value decrease with the rising of reduced frequencies. It is predicted that when the reduced frequency reaches a critical value, depending on the initial angle-ofattack, the hypersonic inlet would keep start/unstart in the whole circle.

CFD simulations were performed on four 2-D hypersonic inlets with different aspect ratios in constant area.Effects of aspect ratio were analyzed at designed Mach 6 and other Mach numbers with different back pressures.The investigation suggest that with the reducing of length and wetted area while rising of aspect ratio,total pressure recovery coefficient of external compress ramp,mass flow ratio and coefficient of internal drag decrease;Inlet with low aspect ratio,as the huge area of three-dimensional-flow near the side wall,performance of inlet is affected notable and the starting ability become worse;When the inlets work in peak back pressure with design Mach number,the pressure that the 2-D hypersonic inlets can endure would reduce with the rising of aspect ratios.