基于模糊神经网络的涡结构识别方法研究

研究气动光学传输效应产生的机理是红外成像末制导的共性基础技术之一,基于涡结构对光学传输效应进行建模是一种非常有效的方法,而涡结构的识别是其必要前提。文中提出一种新的涡结构识别方法,把折射率场经小波变换后的系数矩阵等效为具有一定纹理结构的图像,计算图像的共生矩阵及其统计量,由于涡结构模式复杂,特征量较多,设计了等价结构的模糊神经网络进行涡结构识别。与小波分解后直接提取特征量的识别方法相比,本文的方法从空、频角度更加准确全面地表征湍流涡结构模式,计算机仿真结果表明该方法优于神经网络的识别效率。     

High-resolution simulation for rotorcraft aerodynamics in hovering and vertical descending flight using a hybrid method

A high-resolution simulation tool for rotorcraft aerodynamics is developed by coupling CFD with a Vorticity Transport Model (VTM). An Eulerian-based CFD module is used to model the blade near body flowfield, and a Lagrangian-based VTM module is employed for vortex tracking in the far wake. The coupling procedure is implemented by transmitting vortex sources to the VTM module and feeding boundary conditions back to the CFD module. The presented CFD/VTM hybrid solver is firstly validated by hover cases of three different rotor configurations. Simulation results, including the blade surface pressure distribution, rotor downwash, and hover figure of merit, exhibit favorable correlations with available experimental data. Then, a rotor operated in vertical descending flight with a fixed collective pitch is investigated. It is shown that the CFD/VTM coupling method is suitable for rotor wake simulation. Wake instabilities (far wake breakdown in hover and toroidal wake pattern in the vortex ring state) are successfully demonstrated with a moderate computational cost.     

Numerical investigation on flow field characteristics of dual synthetic cold/hot jets using POD and DMD methods

In order to improve the control ability of synthetic jets in compressible boundary layer, a novel control method based on dual synthetic cold/hot jets coupled control of velocity profile and temperature profile was proposed. As fundamental investigations on the effects of synthetic jet temperature on the jet behavior and flow field characteristics were essentially necessary, preliminary numerical simulations were conducted to study the influence of temperature (200 K and 400 K) on the flow field characteristics of synthetic jets using Large Eddy Simulations (LES) model. Time-averaged flow fields showed that different temperatures led to variable behavior of two strands of jets. For dual synthetic cold jets, a potential-core arose apparently with its height ranging from 0.01 to 0.03 m, while for dual synthetic hot jets, two strands of jets emerged downstream. The modal decomposition of instantaneous flow fields had been done using both Proper Orthogonal Decomposition (POD) and Dynamic Mode Decomposition (DMD). Various modes showed different characteristics of the flow fields. As the POD method focuses on the energy of flow while the DMD method focuses on the frequency, the first two modes had many similarities, but the third and fourth modes demonstrated completely different vortex structures. The current researches play a role of preliminary investigations for further and comprehensive exploration of novel flow control measures in global velocity field.     

Supersonic flow over a pitching delta wing using surface pressure measurements and numerical simulations

Experimental and numerical methods were applied to investigating high subsonic and supersonic flows over a 60° swept delta wing in fixed state and pitching oscillation.Static pressure coefficient distributions over the wing leeward surface and the hysteresis loops of pressure coefficient versus angle of attack at the sensor locations were obtained by wind tunnel tests.Similar results were obtained by numerical simulations which agreed well with the experiments.Flow structure around the wing was also demonstrated by the numerical simulation.Effects of Mach number and angle of attack on pressure distribution curves in static tests were investigated.Effects of various oscillation parameters including Mach number, mean angle of attack, pitching amplitude and frequency on hysteresis loops were investigated in dynamic tests and the associated physical mechanisms were discussed.Vortex breakdown phenomenon over the wing was identified at high angles of attack using the pressure coefficient curves and hysteresis loops, and its effects on the flow features were discussed.     

掺混孔结构对大曲率受限空间出口温度分布的影响

回流燃烧室的掺混射流在大曲率受限的弯管中与主流掺混,回流导致沿径向的压力梯度变化,增加了掺混流动的复杂性.为了研究不同掺混孔结构对掺混射流与主流在大曲率弯曲通道内的掺混规律,在有收缩比弯曲通道中多孔射流的出口温度分布测试中,对影响掺混冷却的各变化参数按照单一变量进行比对研究,对孔径、孔间距等掺混孔结构变量弯曲通道内的掺...     

Experimental investigation on aero-heating of rudder shaft within laminar/turbulent hypersonic boundary layers

The aero-heating of the rudder shaft region of a hypersonic vehicle is very harsh, as the peak heat flux in this region can be even higher than that at the stagnation point. Therefore, studying the aero-heating of the rudder shaft is of great significance for designing the thermal protection system of the hypersonic vehicle. In the wind tunnel test of the aero-heating effect, we find that with the increase of the angle of attack of the lifting body model, the increasement of the heat flux of the rudder shaft is larger under laminar flow conditions than that under turbulent flow conditions. To understand this, we design a wind tunnel experiment to study the effect of laminar/turbulent hypersonic boundary layers on the heat flux of the rudder shaft under the same wind tunnel freestream conditions. The experiment is carried out in the ?2 m shock tunnel(FD-14 A) affiliated to the China Aerodynamics Research and Development Center(CARDC). The laminar boundary layer on the model is triggered to a turbulent one by using vortex generators, which are 2 mm-high diamonds. The aero-heating of the rudder shaft(with the rudder) and the protuberance(without the rudder) are studied in both hypersonic laminar and turbulent boundary layers under the same freestream condition. The nominal Mach numbers are 10 and 12, and the unit Reynolds numbers are2.4 × 10~6 m~(-1) and 2.1 × 10~6 m-1. The angle of attack of the model is 20°, and the deflection angle of the rudder and the protuberance is 10°. The heat flux on the model surface is measured by thin film heat flux sensors, and the heat flux distribution along the center line of the lifting body model suggests that forced transition is achieved in the upstream of the rudder. The test results of the rudder shaft and the protuberance show that the heat flux of the rudder shaft is lower in the turbulent flow than that in the laminar flow, but the heat flux of the protuberance is the other way around,i.e., lower in the laminar flow than in the turbulent flow. The wind tunnel test results is also validated by numerical simulations. Our analysis suggests that this phenomenon is due to the difference of boundary layer velocities caused by different thickness of boundary layer between laminar and turbulent flows, as well as the restricted flow within the rudder gap. When the turbulent boundary layer is more than three times thicker than that of the laminar boundary layer, the heat flux of the rudder shaft under the laminar flow condition is higher than that under the turbulent flow condition. Discovery of this phenomenon has great importance for guiding the design of the thermal protection system for the rudder shaft of hypersonic vehicles.     

短舱涡流发生器流动控制机理初探

首先回顾了短舱涡流发生器在民机研制中的应用及发展历程,然后结合民用飞机研发实践,借助于CFD工具初步计算并着重分析了短舱涡流发生器的流动控制机理。结果分析表明合理设计的短舱涡流发生器能够产生能量较强的集中涡,在分离敏感的短舱后翼面区域产生下洗流场并为边界层注入能量,降低短舱后气流的当地有效迎角,延缓由于"挂架涡"提前破裂而导致的大面积分离,从而能够改善失速区形态。     

三角形埋入式涡流发生器几何参数对其涡的影响

试验研究了三角形埋入式涡流发生器几何参数对其涡各参数的影响，发展了半片式三角形埋入式涡流发生器。试验中采用小型五孔探针获得涡流发生器后不同轴向截面处的速度矢量图，用高斯涡核模型及非线性拟牛顿最小二乘法拟合涡量、涡核半径、涡心高度等参数，分析了这些参数沿程变化趋势。     

超燃冲压发动机燃烧室的积分分析方法

本文以Billig模型和公式为基础,给出了超音燃烧室的分析计算方法;并以燃烧室的进口马赫数是2.2为例,对燃烧室的工况、壁型及参数变化进行了详细的分析研究。所得结论,有助于对超音燃烧室的理解和设计。     

Effects of wing locations on wing rock induced by forebody vortices

Previous studies have shown that asymmetric vortex wakes over slender bodies exhibit a multi-vortex structure with an alternate arrangement along a body axis at high angle of attack. In this investigation, the effects of wing locations along a body axis on wing rock induced by forebody vortices was studied experimentally at a subcritical Reynolds number based on a body diameter. An artificial perturbation was added onto the nose tip to fix the orientations of forebody vortices. Par-ticle image velocimetry was used to identify flow patterns of forebody vortices in static situations, and time histories of wing rock were obtained using a free-to-roll rig. The results show that the wing locations can affect significantly the motion patterns of wing rock owing to the variation of multi-vortex patterns of forebody vortices. As the wing locations make the forebody vortices a two-vortex pattern, the wing body exhibits regularly divergence and fixed-point motion with azimuthal varia-tions of the tip perturbation. If a three-vortex pattern exists over the wing, however, the wing-rock patterns depend on the impact of the highest vortex and newborn vortex. As the three vortices together influence the wing flow, wing-rock patterns exhibit regularly fixed-points and limit-cycled oscillations. With the wing moving backwards, the newborn vortex becomes stronger, and wing-rock patterns become fixed-points, chaotic oscillations, and limit-cycled oscillations. With fur-ther backward movement of wings, the vortices are far away from the upper surface of wings, and the motions exhibit divergence, limit-cycled oscillations and fixed-points. For the rearmost location of the wing, the wing body exhibits stochastic oscillations and fixed-points.