高超声速飞行器表面横缝旋涡结构及气动热环境数值模拟

针对高超声速飞行器表面横缝内部流动,通过求解可压缩Navier-Stokes方程,自主研发了一套能够较好模拟缝隙流动特性的CFD软件,利用该软件对横缝进行了数值模拟。研究表明,对于缝隙内黏性起主导作用的低速流动,通常需要密度足够充分的网格才能获得较好的计算结果。收敛后的壁面热流与文献实验热流吻合较好;缝隙内部旋涡个数近似与缝隙深宽比成比例,与文献中的有关结论完全一致。缝隙壁面热流分布受旋涡结构影响,会出现波浪式变化。因此,合理捕捉旋涡结构对于缝隙壁面热流的数值模拟有着重要意义,本文CFD软件对高超声速飞行器表面横缝的计算结果具有较高可信度。     

Experimental and computational investigation of hybrid formation flight for aerodynamic gain at transonic speed

The influence of the wing-tip vortex of leading aircraft on energy savings, quantified by formation aerodynamic force fraction of the following aircraft, is studied at transonic speed for a matrix of leading aircraft’s vortex locations. The research model adopts the hybrid formation of medium and large aircraft. The leading aircraft is scaled by 2.1%, and the following aircraft is scaled by 1.4%. An aerodynamic benefit “map” is developed to determine the optimum location of the following aircraft relative to the leading aircraft wake and to compare with experimental results, thus validating the use of CFD for the formation flight at cruising speed. The response surface model of aerodynamic gain effect relative to formation parameters is established via numerical calculation and wind tunnel test. The optimal formation parameters and the setting criteria of the study model are optimized. Results show that the wing-tip vortex of large aircraft significantly increases lift and reduces drag on the medium-sized aircraft following it. Reduced drag slightly increases with the flow direction position. With the increase of flow direction distance, the peak area moves from 15% of wing-tip overlap to 20% of overlap. In addition, the maximum drag decreases about 16%, and the maximum lift increases about 12%. The lift drag ratio of the optimal position is increased by 27%, which is twice as large as that of the same scale ratio aircraft formation. Results show that the increase of lift is mainly caused by the increase of suction peak and suction range.     

Instability characteristics of a co-rotating wingtip vortex pair based on bi-global linear stability analysis

The Stereo Particle Image Velocimetry (SPIV) technology is applied to measure the wingtip vortices generated by the up-down symmetrical split winglet. Then, the temporal bi-global Linear Stability Analysis (bi-global LSA) is performed on this nearly equal-strength co-rotating vortex pair, which is composed of an upper vortex (vortex-u) and a down vortex (vortex-d). The results show that the instability eigenvalue spectrum illustrated by (ω_r, ω_i) contains two types of branches: discrete branch and continuous branch. The discrete branch contains the primary branches of vortex-u and vortex-d, the secondary branch of vortex-d and coupled branch, of which all of the eigenvalues are located in the unstable half-plane of ω_i > 0, indicating that the wingtip vortex pair is temporally unstable. By contrast, the eigenvalues of the continuous branch are concentrated on the half-plane of ω_i < 0 and the perturbation modes correspond to the freestream perturbation. In the primary branches of vortex-u and vortex-d, Mode P_u and Mode P_d are the primary perturbation modes, which exhibit the structures enclosed with azimuthal wavenumber m and radial wavenumber n, respectively. Besides, the results of stability curves for vortex-u and vortex-d demonstrate that the instability growth rates of vortex-u are larger than those of vortex-d, and the perturbation energy of Mode P_u is also larger than that of Mode P_d. Moreover, the perturbation energy of Mode P_u is up to 0.02650 and accounts for 33.56% percent in the corresponding branch, thereby indicating that the instability development of wingtip vortex is dominated by Mode P_u. By further investigating the topological structures of Mode P_u and Mode P_d with streamwise wavenumbers, the most unstable perturbation mode with a large azimuthal wavenumber of m = 5–6 is identified, which imposes on the entire core region of vortex-u. This large azimuthal wavenumber perturbation mode can suggest the potential physical-based flow control strategy by manipulating it.     

涡轮叶片尾缘内冷通道旋流冷却特性

针对简化的叶片尾缘,设计了3种旋流冷却结构,即冷气分别从旋流腔中部射流孔、旋流腔异侧射流孔、旋流腔同侧射流孔进出旋流腔,并与常规凸台扰流柱冷却结构进行了对比数值研究,分析其强化换热机理和效果.结果表明:旋流腔的结构和冷气的进流布置对旋流冷却性能的影响很大,冷气从旋流腔某侧射流孔进出的旋流冷却结构不仅在流向截面产生涡旋,在展向截面也会产生涡旋,从而有效强化对流换热;相比凸台扰流柱冷却结构,旋流冷却结构能够增强换热,平均努塞尔数增大6.8%~22.9%,但流动阻力也随之增加;冷气从旋流腔异侧射流孔进出的冷却结构强化换热能力较高;而冷气从旋流腔同侧射流孔进出的冷却结构流动换热综合系数比凸台扰流柱提高4.2%,综合性能相对较优.      

对转开式转子非定常气动干扰特性分析

采用动态面搭接技术,求解非定常雷诺平均Navier-Stokes方程实现对转开式转子的非定常气动数值模拟.对10片桨叶单个转子和10×10对转开式转子构型分别进行模拟,对比分析前后两个转子间的气动干扰和滑流流动干扰.结果表明:与单个转子相比,对转开式转子前转子的拉力系数和功率系数减小,后转子的拉力系数和功率系数则都增大,并在一个旋转周期内都呈现20次周期性波动.前后转子拉力系数的频谱分析显示振荡发生在桨叶通过频率的偶数倍,且在2倍时后转子拉力的振幅最大.前转子桨尖涡与后转子的周期性干扰,引起后转子桨叶力分布的变化和非定常性.与单个转子相比,前转子后滑流轴向速度偏大,周向速度偏小.后转子对滑流有二次加速及旋涡恢复作用.      

基于VVPM/CA耦合的旋翼非定常载荷计算

为了提高旋翼非定常载荷计算精度,将基于黏性涡粒子方法(Viscous vortex particle method,VVPM)的尾迹计算引入至旋翼综合分析(Comprehensive analysis,CA)中,建立了一个新的旋翼VVPM/CA耦合计算模型。该模型中,旋翼VVPM基于第一性原理,可模拟尾迹的畸变和扩散而不引入经验参数,而旋翼CA则可以有效地进行桨叶弹性变形及非定常载荷计算,通过采用松耦合策略,可以高效地实现两者的信息交换。在此基础上,以SA349/2直升机为算例,针对其低速和高速两种典型前飞状态进行了深入分析,计算表明,本文建立的VVPM/CA耦合分析可以有效地预测旋翼尾迹形状及非定常载荷。     

驻涡燃烧室最佳中心驻体宽度选择的数值研究

为确定驻涡燃烧室中心驻体最佳宽度,在四种来流条件下,对当量宽度为0.3～0.8范围,间隔为0.05当量宽度的中心驻体驻涡燃烧室三维黏性流场进行了冷态数值模拟.计算结果表明,当量宽度在0.65～0.70的中心驻体有利于在驻涡腔内形成低速且稳定的流动.驻体宽度增加会导致驻涡腔外侧主流气流速度提高,引起摩擦损失的加大,燃烧室出口截面处的总压损失增加.      

类车体尾迹区流动的实验研究

对后倾角35°类车体在模型风洞中进行实验,运用三维热线风速仪在选定测点上对空间三个速度分量进行采样。对采得的瞬时值进行平均值处理,并制成相应的云图和矢量图。由图谱说明：类车体尾迹区流动具有对称特性,以一对拖曳涡为其主要特征,并明确了拖曳涡的涡核位置。     

翼梢小翼后缘舵面偏转对机翼气动特性影响研究

以大型客机某方案机翼为基本翼,基于N-S方程数值模拟的方法,研究融合式翼梢小翼后缘操纵舵面偏转对机翼空气动力特性影响。研究发现,翼梢小翼舵面偏使得机翼气动特性发生显著变化。一方面,偏转舵面导致了机翼最大升阻比的降低,然而它可以优化不同飞行阶段升阻比。其中,舵面外偏,机翼在阻力增加不大的条件下,升力明显增大,有利于提高起飞、爬升性能;舵面不偏条件下升阻比最大,有利于提高巡航效率;舵面内偏,机翼阻力明显增大,有利于提高飞机着陆性能。另一方面,舵面偏转可以控制机翼翼梢涡的发展,有助于耗散机翼尾涡及激发翼梢涡自身的不稳定性而加速耗散。     

燃气轮机涡轮导向叶片涡流交错肋冷却技术研究

研究了燃气轮机涡轮涡流交错肋冷却导向叶片的冷却效果。采用由联立的连续方程、动量方程和能量方程构成的专用程序,计算了冷却系统流体动力和换热特性;采用通用的有限元计算方法计算了叶片温度场。通过对涡流交错肋通道的几何特性、流体阻力特性参数、叶片冷却系统的流体动力和换热、叶片温度场的计算,获得了气冷导向叶片的冷却效果。考虑到最大径向和周向不均匀度和+30℃裕度的情况,得到气冷叶片的中间截面平均降温为350℃,气冷叶片中截面的平均冷却效果为0.383。