HEAT-PIPE-COOLED THERMAL PROTECTION FOR THE LEADING EDGE OF THE WING

针对高超声速飞行器飞行时翼前缘存在着严重的气动加热问题，为了保证翼前缘的尖锐外形，提出疏导式热防护结构，利用内置高温热管结构为翼前缘提供热防护。采用数值模拟和电弧风洞试验的方法对翼前缘疏导式结构进行了分析，得到翼前缘内置高温热管具有的防热效果。数值模拟结果表明在一定热环境条件下，翼前缘驻点温度下降了304K，尾部最低温度升高了130K，实现了热流从高温区到低温区的疏导，减弱了翼前缘的热载荷，强化了翼前缘的热防护能力。通过电弧风洞试验可以获得相同的热防护结果，并且在一定飞行条件下高温热管可以自适应启动，验证了数值模拟方法的准确性以及翼前缘内置高温热管疏导式热防护结构的可行性。     

Reynolds number effects on leading edge vortex development on a waving wing

The waving wing experiment is a fully three-dimensional simplification of the flapping wing motion observed in nature. The spanwise velocity gradient and wing starting and stopping acceleration that exist on an insect-like flapping wing are generated by rotational motion of a finite span wing. The flow development around a waving wing at Reynolds number between 10,000 and 60,000 has been studied using flow visualization and high-speed PIV to capture the unsteady velocity field. Lift and drag forces have been measured over a range of angles of attack, and the lift curve shape was similar in all cases. A transient high-lift peak approximately 1.5 times the quasi-steady value occurred in the first chord length of travel, caused by the formation of a strong attached leading edge vortex. This vortex appears to develop and shed more quickly at lower Reynolds numbers. The circulation of the leading edge vortex has been measured and agrees well with force data.     

A Numerical Study of the Flow Around a Model Winged Seed in Auto-Rotation

In this study the flow around a winged-seed in auto-rotation is characterized using direct numerical simulations (DNS) at Reynolds number in the range 80–240, based on the descent speed and a characteristic chord length. In this range, the flow is approximately steady when observed from a reference frame fixed to the seed. For all cases, the flow structure consists of a wing tip vortex which describes a helical path, a vortex shed behind the nut of the seed and a stable leading edge vortex above the wing surface which merges with the tip vortex. With increasing Reynolds number, the leading edge vortex becomes more intense and gets closer to the wing surface. The simulation results also show the formation of a spanwise flow on the upper surface of the wing, moving fluid towards the wing tip in a region downstream and beneath the leading edge vortex. This spanwise flow is rather weak inside the core of the leading edge vortex, and the analysis of the streamlines show a very weak transport of vorticity along the vortex for the cases under consideration. The analysis of the flow suggests that the stabilization of the leading edge vortex is mainly due to non-inertial accelerations, although viscous effects may contribute, specially at lower Re. Furthermore, the leading edge vortex has been characterized by analysing the flow variables averaged along cross-sections of the vortex. While some quantities, like the spanwise velocity or the pressure inside the vortex, are rather insensitive to the threshold used to define the leading edge vortex, the same is not true for the circulation of the vortex or its averaged spanwise vorticity, due to the viscous nature of the vortex. Finally, it is observed that the spanwise vorticity scales with the angular rotation of the seed for the different Re.     

基于雨燕翅膀的仿生三角翼气动特性计算研究

针对低雷诺数微型飞行器的气动布局, 设计出类似雨燕翅膀的一组具有不同前缘钝度的中等后掠($\varLambda =50^{\circ}$)仿生三角翼. 为了定量对比研究三角翼后缘收缩产生的气动效应, 设计了一组具有同等后掠的普通三角翼. 为了深入研究仿生三角翼布局的前缘涡演化特性以及总体气动特性, 采用数值模拟方法详细地探索了低雷诺数($Re=1.58\times 10^{4})$流动条件下前缘涡涡流结构和气动力随迎角的变化规律. 分析结果表明, 前缘钝度和后缘收缩对仿生三角翼前缘涡的涡流强度和涡破裂位置有显著影响. 相对于钝前缘来说, 尖前缘使仿生三角翼上下表面的压力差增大, 涡流强度也更大, 增升作用也更显著. 相对于普通三角翼构型, 仿生三角翼的前缘斜切使其阻力更大, 但后缘的收缩使涡破裂位置固定在此位置, 因此整个上翼面保持低压, 总的升力更大. 由于小迎角时升力增大更明显, 因此仿生三角翼的气动效率在小迎角时明显大于普通三角翼. 这些结论对于揭示鸟类的飞行机理以及未来微型仿生飞行器的气动布局设计具有重要的研究价值.      

Combined method of calculating supersonic ideal-gas flow over a wing with a supersonic blunt leading edge

A combined numerical method, based on the successive calculation of the flow regions near the blunt leading edge and center of a wing, is proposed on the assumption that the angle of attack and the relative thickness and bluntness radius of the leading edge are small. The flow in the neighborhood of the leading edge of the wing is assumed to be identical to that on the windward surface of a slender body coinciding in shape with the surface of the blunt nose of the wing and is numerically determined in accordance with [1]. The flow parameters near the center of the wing are calculated within the framework of the law of plane sections [2]. In both regions the equations of motion of the gas are integrated by the Godunov method. The flow fields around elliptic cones are obtained within the framework of the combined method and the method of [3], A comparative analysis of the results of the calculations makes it possible to estimate the region of applicability of the method proposed.Translated from Izvestiya Akademii Nauk SSSR, Mekhanika Zhidkosti i Gaza, No. 1, pp. 159–164, January–February, 1989.The authors wish to express their gratitude to A. A. Gladkov for discussing their work, and to G. P. Voskresenskii, O. V. Ivanov, and V. A. Stebunov for making available a program for calculating supersonic flow over a wing with a detached shock.     

A combined analytical–numerical method for treating corner singularities in viscous flow predictions

A combined analytical–numerical method based on a matching asymptotic algorithm is proposed for treating angular (sharp corner or wedge) singularities in the numerical solution of the Navier–Stokes equations. We adopt an asymptotic solution for the local flow around the angular points based on the Stokes flow approximation and a numerical solution for the global flow outside the singular regions using a finite‐volume method. The coefficients involved in the analytical solution are iteratively updated by matching both solutions in a small region where the Stokes flow approximation holds. Moreover, an error analysis is derived for this method, which serves as a guideline for the practical implementation. The present method is applied to treat the leading‐edge singularity of a semi‐infinite plate. The effect of various influencing factors related to the implementation are evaluated with the help of numerical experiments. The investigation showed that the accuracy of the numerical solution for the flow around the leading edge can be significantly improved with the present method. The results of the numerical experiments support the error analysis and show the desired properties of the new algorithm, i.e. accuracy, robustness and efficiency. Based on the numerical results for the leading‐edge singularity, the validity of various classical approximate models for the flow, such as the Stokes approximation, the inviscid flow model and the boundary layer theory of varying orders are examined. Although the methodology proposed was evaluated for the leading‐edge problem, it is generally applicable to all kinds of angular singularities and all kinds of finite‐discretization methods. Copyright © 2004 John Wiley & Sons, Ltd.     

数值研究激波与旋涡的相互作用

从非定常形式的Euler方程出发,数值模拟了运动激波与旋涡相互作用的非定常流动过程。为保证激波具有较高的分辨率,采用对称型TVD格式进行了数值计算。结果表明。这样可以有效地模拟流场中一些复杂的流动现象,如激波变形、激波分叉和三波点的形成,以及旋涡结构的变化过程等,并与已有的实验流动显示相符良好。同时,也是对TVD格式求解这类问题的一次初步尝试。     

某型飞机机翼前缘抗鸟撞结构设计与试验验证

针对某型飞机机翼前缘抗鸟撞性能不满足适航要求的问题，采用"数值仿真-试验验证-再仿真"的研究思路，对该结构进行了抗鸟撞优化设计。首先，通过有限元数值仿真，分别对具有三角板结构和前墙结构的两种新型前缘结构抗鸟撞能力进行了分析。仿真结果表明:具有前墙结构的机翼前缘抗鸟撞能力优于原始结构和带三角板结构的机翼前缘；在鸟撞过程中，这种具有前墙结构的机翼前缘通过利用破损蒙皮继续变形吸能的方式提高了结构的抗鸟撞性能。基于此，对带前墙结构的机翼前缘进行了抗鸟撞试验，一方面验证了数值模拟方法的准确性，另一方面验证了前墙结构的抗鸟撞效果。最后，采用数值仿真方法对带前墙结构的机翼前缘结构进行了参数分析，得到了前缘蒙皮厚度和前墙厚度与结构抗鸟撞性能的关系，并基于结构承载和抗鸟撞能力的综合要求，确定了最终结构参数。分析表明，优化后的机翼前缘结构不仅满足抗鸟撞要求，而且实现结构减重30%。     

Convergence issues in using high‐resolution schemes and lower–upper symmetric Gauss–Seidel method for steady shock‐induced combustion problems

This paper reports numerical convergence study for simulations of steady shock‐induced combustion problems with high‐resolution shock‐capturing schemes. Five typical schemes are used: the Roe flux‐based monotone upstream‐centered scheme for conservation laws (MUSCL) and weighted essentially non‐oscillatory (WENO) schemes, the Lax–Friedrichs splitting‐based non‐oscillatory no‐free parameter dissipative (NND) and WENO schemes, and the Harten–Yee upwind total variation diminishing (TVD) scheme. These schemes are implemented with the finite volume discretization on structured quadrilateral meshes in dimension‐by‐dimension way and the lower–upper symmetric Gauss–Seidel (LU–SGS) relaxation method for solving the axisymmetric multispecies reactive Navier–Stokes equations. Comparison of iterative convergence between different schemes has been made using supersonic combustion flows around a spherical projectile with Mach numbers M = 3.55 and 6.46 and a ram accelerator with M = 6.7. These test cases were regarded as steady combustion problems in literature. Calculations on gradually refined meshes show that the second‐order NND, MUSCL, and TVD schemes can converge well to steady states from coarse through fine meshes for M = 3.55 case in which shock and combustion fronts are separate, whereas the (nominally) fifth‐order WENO schemes can only converge to some residual level. More interestingly, the numerical results show that all the schemes do not converge to steady‐state solutions for M = 6.46 in the spherical projectile and M = 6.7 in the ram accelerator cases on fine meshes although they all converge on coarser meshes or on fine meshes without chemical reactions. The result is based on the particular preconditioner of LU–SGS scheme. Possible reasons for the nonconvergence in reactive flow simulation are discussed.Copyright © 2012 John Wiley & Sons, Ltd.     

Simulation of transients in natural gas pipelines using hybrid TVD schemes

The mathematical model describing transients in natural gas pipelines constitutes a non‐homogeneous system of non‐linear hyperbolic conservation laws. The time splitting approach is adopted to solve this non‐homogeneous hyperbolic model. At each time step, the non‐homogeneous hyperbolic model is split into a homogeneous hyperbolic model and an ODE operator. An explicit 5‐point, second‐order‐accurate total variation diminishing (TVD) scheme is formulated to solve the homogeneous system of non‐linear hyperbolic conservation laws. Special attention is given to the treatment of boundary conditions at the inlet and the outlet of the pipeline. Hybrid methods involving the Godunov scheme (TVD/Godunov scheme) or the Roe scheme (TVD/Roe scheme) or the Lax–Wendroff scheme (TVD/LW scheme) are used to achieve appropriate boundary handling strategy. A severe condition involving instantaneous closure of a downstream valve is used to test the efficacy of the new schemes. The results produced by the TVD/Roe and TVD/Godunov schemes are excellent and comparable with each other, while the TVD/LW scheme performs reasonably well. The TVD/Roe scheme is applied to simulate the transport of a fast transient in a short pipe and the propagation of a slow transient in a long transmission pipeline. For the first example, the scheme produces excellent results, which capture and maintain the integrity of the wave fronts even after a long time. For the second example, comparisons of computational results are made using different discretizing parameters. Copyright © 2000 John Wiley & Sons, Ltd.     

N-S方程数值研究翼型对微型扑翼气动特性的影响

首先基于嵌套网格发展了一套适用于三维扑翼研究的非定常雷诺平均Navier-Stokes(RANS)方程数值模拟方法.为了解决微型扑翼在低马赫数下的收敛问题,使用了预处理方法,湍流模型为BL模型.在该方法的基础上,保持状态参数和扑翼表面形状一定的情况下,分别研究了一系列不同厚度、不同弯度的翼型对于微型扑翼气动特性的影响....     

Time-resolved scanning tomography PIV measurements around a flapping wing

The three-dimensional flow that develops around a finite flapping wing is investigated using a tomographic scanning PIV technique. The acquisition and correlation processes employed to achieve such measurements have been carefully validated. Issues regarding the relevant timescales of the flow and the spanwise space-resolution are addressed. Results obtained on a hovering flapping wing whose plunging phase is described by a rectilinear motion highlight the influence of the free end condition and the formation of the tip vortex on the leading edge vortices behavior, wing/wake interactions, and wake stabilization.     

Aerodynamic improvement of a delta wing in combination with leading edge flaps

Recently, various studies of micro air vehicle(MAV) and unmanned air vehicle(UAV) have been reported from wide range points of view. The aim of this study is to research the aerodynamic improvement of delta wing in low Reynold's number region to develop an applicative these air vehicle. As an attractive tool in delta wing, leading edge flap(LEF) is employed to directly modify the strength and structure of vortices originating from the separation point along the leading edge. Various configurations of LEF such as drooping apex flap and upward deflected flap are used in combination to enhance the aerodynamic characteristics in the delta wing. The fluid force measurement by six component load cell and particle image velocimetry(PIV) analysis are performed as the experimental method. The relations between the aerodynamic superiority and the vortex behavior around the models are demonstrated.     

Multiple laminar-turbulent transition cycles around a swept leading edge

Certain interesting flow features involving multiple transition/relaminarization cycles on the leading edge of a swept wing at low speeds are reported here. The wing geometry tested had a circular nose and a leading edge sweep of 60°. Tests were made at a chord Reynolds number of 1.3 × 10⁶ with model incidence α varied in the range of 3°?18° in discrete steps. Measurements made included wing chord-wise surface pressure distributions and wall shear stress fluctuations (using hot-film gages) within about 10 % of the chord in the leading edge zone. Results at α = 16° and 18° showed that several (often incomplete) transition cycles between laminar-like and turbulent-like flows occurred. These rather surprising results are attributable chiefly to the fact that the Launder acceleration parameter K (appropriately modified for swept wings) can exceed a critical range more than once along the contour of the airfoil in the leading edge region. Each such crossing results in a relaminarization followed by direct retransition to turbulence as K drops to sufficiently low values. It is further shown that the extent of each observed transition zone (of either type) is consistent with earlier data acquired in more detailed studies of direct transition and relaminarization. Swept leading edge boundary layers therefore pose strong challenges to numerical modelling.     

Vortex breakdown and its control on delta wings

In order to investigate the breakdown of vortices generated by the leading edge of delta wings, LDA-measurements have been performed in the flow on the suction side of a delta wing of aspect ratio A = 2. The measurements describe the growth of the vortex along the leading edge and reveal a certain radial structure upstream of the breakdown point. Moreover they shed light on the mechanism responsible for the onset of vortex breakdown on the suction side of a wing.

The occurrence of the breakdown phenomenon on a delta wing may be prevented or at least retarded by the use of spanwise blowing jets. The interaction of vortex and jets giving rise to these effects will also be discussed with the help of measured velocity profiles.     

The numerical simulations of explosion and implosion in air: use of a modified Harten's TVD scheme

Numerical simulations of explosion and implosion in air are carried out with a modified Harten's TVD scheme. The new scheme has a high resolution for contact discontinuities in addition to maintaining the good features of Harten's TVD scheme. In the numerical experiment of spherical explosion in air, the second shock wave (which does not exist in the one‐dimensional shock tube problem) and its subsequent implosion on the origin have been successfully captured. The positions of the main shock wave, the contact discontinuity and the second shock wave have shown satisfactory agreement with those predicted from previous analysis. The numerical results are also compared with those obtained experimentally. Finally, simulations of a cylindrical explosion and implosion in air are carried out. Results of the cylindrical implosion in air are compared with those of previous work, including the interaction of the reflected main shock wave with the contact discontinuity and the formation of a second shock wave. All these attest to the successful use of the modified Harten's TVD scheme for the simulations of shock waves arising from explosion and implosion. Copyright © 1999 John Wiley & Sons, Ltd.     

连续分层海洋中内波传播的一种数值模式

基于Euler方程,使用有限体积法建立了一种密度为连续分层情况下、适应水深变化的水域中内波传播的数值模式.为了使计算格式能够达到二阶精度,对流项的处理使用了TVD (total variation diminishing)格式.将SIMPLE算法引入连续分层海洋中内波的数值计算,为了简化计算并方便地适应多种TVD格式,在计算预估速度场时采用了显式格式,而没有采用传统的隐式格式;鉴于在原始的SIMPLE算法中没有涉及到由于密度扰动而引起的静水压力场的改变问题,给出了该问题的计算方法.因此改进了SIMPLE算法.出流边界的处理采用阻尼消波和Sommerfeld辐射条件相结合的方式,以使内波得到有效的衰减和释放.将等水深水域的数值解和理论解进行了比较,两者吻合较好;并对存在潜堤时数值计算的不同时刻密度变化的空间分布进行了详细的定性分析.计算结果表明,所建立的数值模式能有效地模拟内波的传播和变形.      

近水面水下爆炸二维Level-set数值模拟

水下爆炸是一个多介质、动边界、瞬态非线性过程,捕捉介质之间交界面的变化是数值模拟的一个难点.本文采用Level-set方法来描述水下爆炸气体产物和流体交界面以及自由表面变化,应用TVD (Total Variation Diminishing) 计算技术求解Level-set流场,首先建立了近水面水下爆炸的二维数值方法,然后,将该方法推广到流场内有刚性结构的情况,考虑了冲击波和刚性结构物的相互作用.为了模拟无限边界流场问题,引进了Thompson建立的无反射边界条件.     

Vortex control by the spanwise suction flow on the upper surface of delta wing

The numerical investigation has been performed to explore the feasibility of vortex control by leading edge sucking excitation on a delta wing. The results reveal that the flow on the upper surface of the delta wing changes significantly in a wide range of the angle of attack. For the vortical flow at moderate angle of attack, the secondary and tertiary vortices are weakened or suppressed, and the total lift is almost unchanged. For the stalled flow at high angle of attack, the leading edge concentrated vortex is recovered, and the lift is enhanced with increasing suction rate. For the bluff-body flow at even high angles of attack, the lift can still be improved. The concentrated vortex disappears on the upper surface, and the load increment is nearly unchanged along the chordwise direction. The project supported by the National Natural Science Foundation of China (19802018).     

基于三路舵机驱动的扑翼飞行模型气动特性研究

扑动而形成非定常气动现象是扑翼飞行过程中产生高升力的主要原因。本文以Ellington实验的鹰蛾翅膀为原形,设计扑翼实验及数值计算模型。通过压差传感器对翅膀模型上翼面固定位置进行测压,分析前缘涡的产生及脱落情况(考虑动压效应)。测量上下翼面固定位置处的压差,揭示扑翼飞行中产生高升力的主要原因。利用烟风洞观察扑翼模型周围流场结构及特殊涡产生变化情况。另外,根据Ellington提供的升力关系式估算了扑翼模型在一个周期内的平均升力。最后,基于三维欧拉方程对扑翼飞行气动特性进行数值模拟,计算结果与实验吻合良好。