多旋流器阵列贫油直喷燃烧室流场的数值模拟

采用雷诺应力模型(RSM)对于同旋向多旋流器阵列贫油直喷(LDI)燃烧室流场进行了数值模拟.计算结果表明,9个旋流器产生的回流区迅速变形衰减,在远离旋流器阵列的下游,形成了一个大的总体旋流.流场中存在的多个回流区及强湍流表明该燃烧室有潜力实现良好的燃烧性能.通过将RSM的模拟结果与激光多普勒测速仪(LDV)测量值进行比较发现,RSM准确描述了绝大部分流动特性,较好地求解了回流区和高速度梯度问题.      

贫燃料湍流预混燃烧火焰稳定性的数值研究

对突扩燃烧室内甲烷-空气预混燃烧进行了数值模拟,时均控制方程组的封闭采用RNG 湍流输运模型和premixed combustion模型.通过模拟研究了不同当量比及进口速度对甲烷-空气预混燃烧效果的影响.CFD数值模拟结果给出了燃烧室内湍流预混反应流的速度场及温度场分布.模拟结果表明当量比为0.6、进口速度为30 m/s时燃烧室内产生稳定预混火焰.该结果对于航空燃气轮机低NOx排放燃烧技术的研究提供了一定的参考条件.     

燃烧室贫油熄火极限数值预测

在任意曲线坐标系下分别对双级轴向涡流器矩形燃烧室及涡流杯回流环形燃烧室三维两相反应流场进行数值模拟,采用RNG(renormalization group)k-ε,SOM-EBU(second order moment-eddy break up model)及六通量辐射等模型考虑紊流流动、化学反应及辐射传热对燃烧过程的影响;采用SI MPLE(semi-implicit method for pressure linked equation)算法求解流场,气液两相耦合采用颗粒源项法处理.采用提出的燃油稳态逐次逼近法对贫油熄火极限进行数值预测,所得计算值与实验数据较为相符,表明此数值预测方法可靠,可用于工程实践.      

使用特征参数准则预测航空发动机燃烧室贫油熄火极限

阐述了特征参数准则的核心原理和预测流程,使用计算流体力学软件FLUENT数值模拟分析了某型航空发动机环形燃烧室的贫油熄火过程,提出了判断熄火的征兆——"M"型火焰及其产生原因.运用特征参数准则研究了不同进口气流速度、气流温度等关键参数对航空发动机燃烧室贫油熄火极限的作用规律.结果表明:贫油熄火油气比随进口气流速度的增加逐渐下降;当进口气流温度小于395K时,贫油熄火油气比随进口气流温度的增加逐渐降低;当进口气流温度超过395K后,贫油熄火极限基本不随进口气流温度的变化而改变.     

模型燃烧室低频不稳定燃烧的初步探讨

对模型燃烧室在贫油预混条件下产生的燃烧不稳定性的特性进行了初步研究。实验研究了当量比、预混气的喷射速度和燃烧室出口面积对不稳定燃烧的频率和幅值的影响。实验结果表明:在很宽的油气比和速度范围内,出现了燃烧不稳定现象,但振荡的主频集中在200～235Hz之间;随着预混气喷射速度的增加,燃烧不稳定发生的频率和幅值都增加;随着燃烧室出口面积的减小,不稳定燃烧发生的振荡幅值增加而频率却减小,表明不稳定模态发生了变化。      

基于单扇区燃烧室试验的热声模拟方法研究及验证

利用自主开发的热声网络模型研究了不同试验工况、出口边界条件、入口测量段长度对单扇区燃烧室试验台燃烧不稳定性的影响,并与试验结果进行了对比分析.结果表明:声学网络模型准确预测了不同试验工况的失稳频率.预测失稳频率与试验失稳频率相差±10 Hz以内,误差在2％以内.前期单扇区燃烧室试验表明,当入口测量段为0.4 m时,发生...     

三旋流器头部燃烧室拓宽燃烧稳定工作范围的研究

对有三旋流器的单头部燃烧室进行了燃烧性能的实验研究,进行了慢车工况的贫油熄火过程及大工况的冒烟性能的研究。研究的因素考虑了不同的头部油气比(余气系数)及喷嘴形式(单或双油路)。研究结果表明:头部余气系数小时改善了贫油熄火油气比;同时,三旋流器与双油路离心喷嘴组合方案的试验表明,这种方案基本上能满足拓宽燃烧稳定工作范围的要求,但其冒烟与要求相比,其性能已与要求值相接近,在这方面尚需进一步改善。      

单头部/扇形/全环燃烧室贫油点火性能换算

通过单头部、扇形和全环燃烧室的点火性能试验，对比了3种试验件贫油点火边界结果的关联与差异，同时对差异原因进行了讨论分析。进一步地，基于Lefebvre点火模型，通过数据拟合处理方法获得了各试验件贫油点火边界的经验公式，并得到它们之间的换算方程。研究表明:不同类型燃烧室试验件的贫油点火边界变化规律相同；同一状态做对比，全环试验件的贫油点火油气比则明显低于单头部和扇形件的相应值。      

Non-synchronous blade vibration analysis of a transonic fan

This study numerically investigates the aeromechanic behavior of a transonic fan model with a flat tip-leading-edge on the NASA rotor 67 test case. Single-passage unsteady calculations at a near stall operating point of 82% design speed show that the dominant frequencies of mass flow were not the harmonics of the rotor rotational frequency. A full-annulus fluid–structure interaction analysis was subsequently carried out to examine the unsteady flows and their interactions with blade vibrations. The results show that the modal displacement of the backward traveling seventh nodal diameter of the second torsion mode grew exponentially, which reveals that the blade vibration was non-synchronous. The vibration pattern indicates that the aerodynamic mode was resonant with the structural vibration mode. Around the rotor tip, the circumferential vortical propagation induced by interactions among the main flow, tip leakage flow, and tip clearance vortex was the source of aerodynamic excitation. To clarify the mechanism of the non-synchronous vibration, the coupling between aerodynamic disturbance and structural response, i.e., aliasing, was summarized. The frequency spectra of the fluctuating pressure show that an aerodynamic Backward Traveling Wave (BTW) was co-aliased to a structural BTW due to the propagation of the circumferential vortex. The correlation between the frequency and free convective speed of the aerodynamic disturbance determined the directions of aliasing.     

Adaptive feedback control of stability in an axial compressor with circumferential inlet distortion

The adaptive feedback control of stability with circumferential inlet distortion has been experimentally investigated in a low-speed, axial compressor. The flat-baffles with different span heights are used to simulate different distorted inflow cases. Compared with auto-correlation and root-mean-square analysis, cross-correlation analysis used to predict early stall warning does not depend on the distortion position. Hence, the cross-correlation coefficient was used to monitor the stable status of the compressor and provide the feedback signal in the active control strategy when suffering from different distortions. Based on the stall margin improvement of tip air injection obtained under different distorted inflow cases and the sensitivity analysis of cross-correlation coefficients to injected momentum ratios, tip air injection was adopted as the actuator for adaptive feedback control. The digital signal processing controller was designed and applied to achieve adaptive feedback control in distorted inflow conditions. The results show that the adaptive feedback control of air injection nearly achieves the same stall margin improvement as steady air injection under different distortion intensities with a reduced injection mass flow. Thus, the proposed adaptive feedback control method is ideal for the engine operation with circumferential distorted inflow, which frequently occurs in flight.