某型电动飞机气动特性数值模拟与风洞实验

为了验证某型大展弦比电动飞机气动设计的合理性以及为飞行性能及品质计算提供数据,采用有限体积法离散求解三维可压雷诺平均Navier-Stokes方程,并选用Spalart-Allmaras湍流模型对该电动飞机流场进行CFD数值模拟。结果表明,该型电动飞机气动设计合理,巡航速度升阻比最高可达23,具有较高气动效率;通过CFD数值模拟得到了全机升力系数、阻力系数和升阻比。为了验证CFD计算结果,对该型电动飞机进行了缩比模型风洞实验,结果显示,CFD数值模拟法计算结果与风洞实验结果高度吻合,说明CFD计算结果准确。该方法可为大展弦比电动飞机气动设计提供指导。     

GALLOPING BEHAVIORS OF SECTOR-SHAPE ICED EIGHT BUNDLE CONDUCTORS1)

考虑到重冰区的特高压输电线路更容易形成接近扇形的覆冰,结合风洞实验及数值模拟方法研究了扇形覆冰八分裂导线的舞动特征。通过扇形覆冰八分裂输电线的节段模型风洞实验,获得各覆冰子导线随风攻角变化的气动参数。在ABAQUS软件中建立单档扇形覆冰八分裂输电线路的有限元模型,通过程序UEL输入各覆冰子导线的气动参数,根据数值模拟结果可获得舞动轨迹及振幅等舞动特征,最后讨论了风速、档距和初始风攻角对八分裂导线舞动特征的影响。     

扇形覆冰特高压八分裂导线舞动特性分析

考虑到重冰区的特高压输电线路更容易形成接近扇形的覆冰,结合风洞实验及数值模拟方法研究了扇形覆冰八分裂导线的舞动特征。通过扇形覆冰八分裂输电线的节段模型风洞实验,获得各覆冰子导线随风攻角变化的气动参数。在ABAQUS软件中建立单档扇形覆冰八分裂输电线路的有限元模型,通过程序UEL输入各覆冰子导线的气动参数,根据数值模拟结果可获得舞动轨迹及振幅等舞动特征,最后讨论了风速、档距和初始风攻角对八分裂导线舞动特征的影响。     

高超声速飞行器气动热关联换算方法研究

气动热风洞实验是地面研究和预测飞行器气动热环境的重要手段之一, 但由于风洞实验模拟能力的限制, 风洞实验的流场参数和模型的几何尺度都会与实际飞行情况存在一定的差别, 导致地面风洞实验中得到的模型表面气动加热率数据无法直接用于飞行条件下的热环境预测和热防护设计. 以往通过针对具体飞行器的试验结果进行数据拟合后外插的气动热关联换算方法指向性较强, 没有考虑到气动热的具体影响参数, 存在一定局限性, 难以外推应用于其他外形的飞行器. 为解决通过气动热风洞实验数据外推预测飞行条件下气动热的技术难题, 基于无量纲NS方程和边界层理论分析研究了影响气动热的主要参数, 并通过推导化简边界层近似解热流公式, 针对层流流态建立了气动热关联换算方法, 可以考虑当地边界层外缘参数的影响, 具有一定通用性. 在此基础上, 利用建立的方法将Reentry-F飞行器缩比模型的风洞实验数据换算到该飞行器飞行条件下的典型工况, 并与飞行测量结果进行了比较, 外推预测结果与飞行测量结果符合较好, 表明建立的关联方法可以用于气动热风洞实验数据的外推换算.      

连拱式大跨悬挑屋盖抗风气动措施研究

本文对连拱式大跨度悬挑屋盖进行了数值风洞试验,分别探讨了在屋盖悬挑前缘增设通风孔、导流板以及同时布设导流板与通风孔这三种气动措施对悬挑屋盖表面风荷载的影响及其作用机理．研究结果表明,同时布设导流板与通风孔的综合气动措施,能显著影响屋盖结构前缘气流的分离,从而减小屋盖的表面风压,并可减弱屋盖的风致振动．在此类屋盖结构设计中,运用综合气动措施可有效降低屋盖整体的升力系数与弯矩系数,对该类屋盖抗风设计较为有利．     

高超声速飞行器气动热关联换算方法研究

气动热风洞实验是地面研究和预测飞行器气动热环境的重要手段之一,但由于风洞实验模拟能力的限制,风洞实验的流场参数和模型的几何尺度都会与实际飞行情况存在一定的差别,导致地面风洞实验中得到的模型表面气动加热率数据无法直接用于飞行条件下的热环境预测和热防护设计.以往通过针对具体飞行器的试验结果进行数据拟合后外插的气动热关联换算方法指向性较强,没有考虑到气动热的具体影响参数,存在一定局限性,难以外推应用于其他外形的飞行器.为解决通过气动热风洞实验数据外推预测飞行条件下气动热的技术难题,基于无量纲NS方程和边界层理论分析研究了影响气动热的主要参数,并通过推导化简边界层近似解热流公式,针对层流流态建立了气动热关联换算方法,可以考虑当地边界层外缘参数的影响,具有一定通用性.在此基础上,利用建立的方法将Reentry-F飞行器缩比模型的风洞实验数据换算到该飞行器飞行条件下的典型工况,并与飞行测量结果进行了比较,外推预测结果与飞行测量结果符合较好,表明建立的关联方法可以用于气动热风洞实验数据的外推换算.     

激波风洞高低压段钢膜片破裂特性研究

激波风洞是用于高超声速飞行器气动外形设计和优化的常用地面试验装置,基于爆轰驱动技术,激波风洞能够在短时间(毫秒级)内产生高温、高压的驱动气体来模拟高超声速试验气流.主膜片位于激波风洞中的爆轰驱动段和激波管段之间,试验时膜片在爆轰脉冲压力下打开,膜片的打开状态和脱落情况对激波风洞气流品质有很大的影响. 同时,膜片也是形成激波的先决条件. 传统的风洞采用铝质膜片进行试验,在激波风洞中需要承压能力更强的膜片, 此时铝质膜片不再适用, 需要采用钢质膜片.因此, 对激波风洞中的钢膜片破裂特性进行研究很有必要.将数值计算结果与试验结果进行比较, 发现数值计算结果与试验结果吻合得比较理想,计算结果具有可靠性. 基于膜片的应力-应变模型, 建立了膜片打开的动力学模型,根据CJ爆轰理论, 采用有限元软件计算模拟了膜片破裂的过程,分析总结了膜片破裂的机制和力学特性规律.采用控制变量法对不同厚度和凹槽长度的膜片进行分析研究,得到了膜片破膜压力和有效破膜时间的变化规律. 在激波风洞试验中,根据膜片总破膜时间设计了适用于JF-12复现风洞的膜片参数.      

SWT-120风洞稳定段的性能测量

在普通超音速风洞中,由于受到噪声干扰,很难进行有效的边界层转捩特性试验研究,针对超声速流动特点发展较低噪声风洞十分必要,而稳定段设计的好坏直接影响到下游试验段噪声水平。本文介绍一座低噪声风洞稳定段的结构和性能测量结果。先对稳定段的结构设计做了简单介绍,然后对实验结果进行分析,实验结果表明在大角度扩散段内装置孔锥,稳定段安装消音夹层和阻尼网组等部件后,气流的速度脉动和压力脉动明显降低,其中压力脉动降低一个量级,速度脉动为1%。进一步优化设计和改进工艺,速度脉动还可进一步降低。测量结果表明SWT-120稳定段的设计是成功的,对我们以后发展更高性能的静风洞有借鉴和参考价值。     

孤立两叶螺旋桨风洞试验准定常数值模拟

为研究某型两叶螺旋桨的气动性能,对某型两叶螺旋桨进行了风洞试验及计算流体动力学(CFD)数值模拟。通过求解多重参考坐标系(MRF)模型下的准定常雷诺平均Navier-Stokes方程,计算了直径为960mm的某型两叶木质螺旋桨静态和动态的气动特性。并在西北工业大学NF-3风洞的三元试验段获得了该螺旋桨桨叶的拉力、扭矩、功率等气动性能数据。螺旋桨风洞试验与CFD数值模拟对比验证表明:基于MRF模型的数值模拟结果与风洞试验结果具有一致性,两者的螺旋桨拉力的偏差控制在5%以内,扭矩的偏差控制在10%以内,拉力功率比的偏差控制在6.08%以内,证明了CFD数值模拟对螺旋桨气动性能预测的准确性。本文可为通用电动飞机螺旋桨的设计与模拟验证提供参考。     

孔板消减气流脉动的数值模拟及实验研究

添加孔板是一种消减压缩机管道系统内气流脉动有效而简便的方法,尽管在工业生产中已被广泛应用,但是其设计和制作所需的各个参数尚处于靠经验取值的阶段.针对这种情况,首先阐述了孔板消减管道内气流脉动的机理;然后使用流体仿真计算软件Fluent建立了管道内气体的二维非稳定流动模型,计算了孔板对管道内气流的压力脉动的影响;并在数值模拟的基础上,搭建了往复式压缩机管道系统实验平台,在进气管线研究了孔板对气流脉动的消减作用.通过数值模拟和实验研究分析了孔板孔径比对气流脉动的影响,并指出选用恰当孔径比的孔板不仅能有效降低主管线和缓冲器至孔板段管线的压力脉动幅度而且对压缩机进口段管线内压力脉动同样具有良好的消减效果.     

Effects of swirl on flow separation and performance of wide angle diffusers

Experimental investigations have been carried out to determine whether the introduction of a circumferential velocity component can produce worthwhile improvements in the performance of, and eliminate flow separation in, wide angle conical diffusers. The swirl generator is a 24 flat-bladed, radial intake type. Systematic experimentation has been carried out for one diffuser configuration fitted with a tailpipe (16.5° and 4.4 area ratio) using varying strengths of inlet swirl and introducing the dissipated mechanical energy as the main criterion of diffuser performance. The best inlet swirl strength produced about 60% reduction of the total diffuser losses in swirl-free flow. The analysis of these results, together with information obtained from flow visualisation experiments, suggests that increasing the swirl beyond an observed threshold completely eliminated flow separation, but it also gave rise to a central zone of recirculating flow and hence additional dissipative losses. We conclude that the optimum improvement achievable in wide angle diffuser performance using swirl does not require the addition of more energy than it saves     

Turbulent flow in a conical diffuser: A review

This is a review of experimental studies of turbulent flow in a conical diffuser by eight Ph.D. students, eleven M.Sc. students, one M.Eng. student, and myself in the past 29 years. During this time, two conical diffusers were constructed: the first was of cast aluminum construction, and the second was of plastic fabrication. These two diffusers were basically the same in geometry except that the pipe section was constructed as an integral part of the plastic diffuser to avoid the lip at the junction of the inlet pipe and the diffuser. The conical diffuser had a total divergence angle of 8°, an area ratio of 4:1, and an inlet diameter of 0.1016 m (4 in.).

The flow at the inlet of the diffuser was usually fully developed pipe flow, but sometimes it was boundary layer grown on the pipe wall. Hot-wire and pulse-wire anemometry together with computer facilities were used to obtain the results of complex flow present in the conical diffuser. Mean velocity profiles were obtained throughout the diffuser, which in turn were used to obtain strain rates and their principal direction. Turbulence moments up to fourth order were measured. The results were used to assess momentum, turbulent kinetic energy, and shear stress equations. Other features such as instantaneous flow reversals in the wall region, relative strength of large eddies, extra strain rate, and the production of kinetic energy also were investigated to find the dynamical picture in the diffuser flow.     

小型高速离心压气机级内部的三维流场分析

利用三维数值模拟技术对微型燃气轮机中的离心压气机部分进行了数值分析,得到了离心压气机设计转速下的级特性曲线和各通流部件中的流动情况。数值分析表明:设计转速下压气机的级特性非常陡峭;整个特性线范围内离心叶轮基本在亚音速情况下工作,而径向扩压器是在跨音速条件下工作,离心压气机整机的最大流量是由径向扩压器的喉部面积决定的;离心压气机级内部各通流部件之间流动的相互干扰是引起流动分离的重要原因,各通流部件之间流动的相互匹配和协调将决定了离心压气机整机的性能和稳定性。     

Experimental and Numerical Investigation of Flows in Complex Shaped Axisymmetric Channels with Mass Injection

The results of an experimental and numerical investigation of the flow in an axisymmetric channel with radial cavities are presented. The experiments were performed with air in a wind tunnel. The theoretical results obtained by numerically solving the problem of viscous axisymmetric gas flow agree with the experimental data on the longitudinal velocity profiles in a channel with a radial cavity.     

Simulations of Turbulent Flow in a Plane Asymmetric Diffuser

Large-eddy simulations (LES) of a planar, asymmetric diffuser flow have been performed. The diverging angle of the inclined wall of the diffuser is chosen as 8.5°, a case for which recent experimental data are available. Reasonable agreement between the LES and the experiments is obtained. The numerical method is further validated for diffuser flow with the diffuser wall inclined at a diverging angle of 10°, which has served as a test case for a number of experimental as well as numerical studies in the literature (LES, RANS). For the present results, the subgrid-scale stresses have been closed using the dynamic Smagorinsky model. A resolution study has been performed, highlighting the disparity of the relevant temporal and spatial scales and thus the sensitivity of the simulation results to the specific numerical grids used. The effect of different Reynolds numbers of the inflowing, fully turbulent channel flow has been studied, in particular, Re _b  = 4,500, Re _b  = 9,000 and Re _b  = 20,000 with Re _b being the Reynolds number based on the bulk velocity and channel half width. The results consistently show that by increasing the Reynolds number a clear trend towards a larger separated region is evident; at least for the studied, comparably low Reynolds-number regime. It is further shown that the small separated region occurring at the diffuser throat shows the opposite behaviour as the main separation region, i.e. the flow is separating less with higher Re _b . Moreover, the influence of the Reynolds number on the internal layer occurring at the non-inclined wall described in a recent study has also been assessed. It can be concluded that this region close to the upper, straight wall, is more distinct for larger Re _b . Additionally, the influence of temporal correlations arising from the commonly used periodic turbulent channel flow as inflow condition (similar to a precursor simulation) for the diffuser is assessed.     

Turbulent separated reattached flow in a two-dimensional curved-wall diffuser

A turbulent separation-reattachment flow in a two-dimensional asymmetrical curved-wall diffuser is studied by a two-dimensional laser doppler velocimeter. The turbulent boundary layer separates on the lower curved wall under strong pressure gradient and then reattaches on a parallel channel. At the inlet of the diffuser, Reynolds number based on the diffuser height is 1.2×10⁵ and the velocity is 25.2m/s. The results of experiments are presented and analyzed in new defined streamline-aligned coordinates. The experiment shows that after Transitory Detachment Reynolds shear stress is negative in the near-wall backflow region. Their characteristics are approximately the same as in simple turbulent shear layers near the maximum Reynolds shear stress. A scale is formed using the maximum Reynolds shear stresses. It is found that a Reynolds shear stress similarity exists from separation to reattachment and the Schofield-Perry velocity law exists in the forward shear flow. Both profiles are used in the experimental work that leads to the design of a new eddy-viscosity model. The length scale is taken from that developed by Schofield and Perry. The composite velocity scale is formed by the maximum Reynolds shear stress and the Schofield-Perry velocity scale as well as the edge velocity of the boundary layer. The results of these experiments are presented in this paper.     

Amplification and amplitude limitation of heave/pitch limit-cycle oscillations close to the transonic dip

Recent results from flutter experiments of the supercritical airfoil NLR 7301 at flow conditions close to the transonic dip are presented. The airfoil was mounted with two degrees-of-freedom in an adaptive solid-wall wind tunnel, and boundary-layer transition was tripped. Flutter boundaries exhibiting a transonic dip were determined and limit-cycle oscillations (LCOs) were measured. The local energy exchange between the fluid and the structure during LCOs is examined and leads to the following findings: at supercritical Mach numbers below that of the transonic-dip minimum the presence of a shock-wave and its dynamics destabilizes the aeroelastic system such that the decreasing branch of the transonic dip develops. At higher Mach numbers the shock-wave motion has a stabilizing effect such that the flutter boundary increases to higher flutter-speed indices with increasing Mach number. Amplified oscillations near this branch of the flutter boundary obtain energy from the flow mainly due to the dynamics of a trailing-edge flow separation. A slight nonlinear amplitude dependency of the shock motion and a possibly occurring boundary-layer separation cause the amplitude limitation of the observed LCOs. The impact of the findings on the numerical simulation of these phenomena is discussed.     

Experiments on transonic limit-cycle-flutter of a flexible swept wing

The aeroelastic behavior of wing models is nonlinear particularly in the transonic speed range. The interaction between aerodynamic and structural forces can lead to the occurrence of Limit-Cycle Oscillations (LCOs). If in addition the wing model is flexible and backward swept, the kinematic coupling between bending and torsion makes the situation even more complex.In the research project “Aerostabil” such a wing was investigated, which was equipped with pressure transducers in three sections and accelerometers. The experiments were performed in the adaptive test section of the transonic wind tunnel TWG in Göttingen. Already Dietz et al. (2003) have reported about experimental details and preliminary results. Based on these data Bendiksen (2008) studied numerically LCO-flutter behavior using a very similar, theoretical model (G-wing) and Stickan et al. (2014) used the original data as a LCO flutter test case. The influence of flexibility on the steady aerodynamics of the wing was described in Schewe & Mai (2018). In this paper now the flutter experiments with the same flexible model were analyzed systematically in the transonic range 0.84 <Ma <0.89 and for six angles of attack from 1.46°to 2.7°. Maps of stability, LCO amplitudes and instantaneous pressure distributions are presented. It was found that unstable regions are islands, whose extent depends on the angle of attack. A LCO test case, already treated in the literature is examined in more detail. The analysis of the time functions showed that during LCO-flutter the motion induced aerodynamic sectional lift forces particularly in the outer wing are asymmetric and thus acting as amplitude limiter. The reason for the asymmetry lies in the shock/boundary layer interaction. The test case, containing the stages of built-up and the transition to the limit cycle provides an excellent opportunity for improving our knowledge about LCOs and for code validation purposes.     

LES of the adverse-pressure gradient turbulent boundary layer

We describe large-eddy simulations (LES) of the flat-plate turbulent boundary layer in the presence of an adverse pressure gradient. The stretched-vortex subgrid-scale model is used in the domain of the flow coupled to a wall model that explicitly accounts for the presence of a finite pressure gradient. The LES are designed to match recent experiments conducted at the University of Melbourne wind tunnel where a plate section with zero pressure gradient is followed by section with constant adverse pressure gradient. First, LES are described at Reynolds numbers based on the local free-stream velocity and the local momentum thickness in the range 6560–13,900 chosen to match the experimental conditions. This is followed by a discussion of further LES at Reynolds numbers at approximately 10 times and 100 times these values, which are well out of range of present day direct numerical simulation and wall-resolved LES. For the lower Reynolds number runs, mean velocity profiles, one-point turbulent statistics of the velocity fluctuations, skin friction and the Clauser and acceleration parameters along the streamwise, adverse pressure-gradient domain are compared to the experimental measurements. For the full range of LES, the relationship of the skin-friction coefficient, in the form of the ratio of the local free-stream velocity to the local friction velocity, to both Reynolds number and the Clauser parameter is explored. At large Reynolds numbers, a region of collapse is found that is well described by a simple log-like empirical relationship over two orders of magnitude. This is expected to be useful for constant adverse-pressure gradient flows. It is concluded that the present adverse pressure gradient boundary layers are far from an equilibrium state.