Based on a turbulence model experimental validation, some numerical simulations were conducted to investigate the effect of plug film cooling on cooling and aerodynamic characteristics for a 2-D plug nozzle. In this study, two cooling schemes namely the half-covering and full-covering film cooling were designed and involved. For each cooling configuration, the effects of perforated percentage (ranged from 0.5% to 2.0%) and non-dimensional inlet total pressure (ranged from 1.02 to 1.20) of coolant on the cooling and aerodynamic performances are discussed separately. The results confirm that an obvious reduction of plug temperature is achieved under both cooling schemes. While the full-covering film cooling configuration performs better, especially at the plug trailing edge where shock waves occurred. Within the scope of the investigation, a small perforated percentage of 0.5% is favorable due to its superior ability to improve the pressure difference between mainstream and coolant; inversely, a big perforated percentage leads to the coolant flow resistance increasing, which may cause the mainstream to flow back at the plug trailing edge. Rising the inlet total pressure of coolant could enhance the cooling performance but also deteriorate the aerodynamic performance of nozzle. Specifically, compared with the case without cooling configuration, the non-dimensional inlet total pressure of coolant ranged from 1.02 to 1.20, the surface temperature of plug decreased by 20% - 45%; the total pressure recovery coefficient decreased by 0.22% - 1.26%.
Aiming at the difficulty of extracting fault features on the aircraft landing gear hydraulic system, traditional feature extraction methods rely heavily on expert knowledge, and the accuracy of fault diagnosis is difficult to guarantee. This paper combined convolutional neural network (CNN) and support vector machine classification algorithm (SVM) to propose a fault diagnosis model suitable for aircraft landing gear hydraulic system. The diagnosis model adopted the one-dimensional multi-channel CNN network structure, took the original pressure signal of multiple nodes as input, adaptively extracts the feature value of the pressure signal through CNN, and built a multi-feature fusion layer to realize the feature fusion of the pressure signal of each node. Finally, input the fused features into the SVM classifier to complete the fault classification. In order to verify the proposed fault diagnosis model, a typical aircraft landing gear hydraulic system simulation model was built based on AMESim, and several typical fault types such as hydraulic pump leakage, actuator leakage, selector valve clogging and accumulator failure were simulated, and corresponding Fault type data set, and use overlapping sample segmentation for data enhancement. Experiments show that the diagnosis accuracy of the proposed fault diagnosis algorithm can reach 99.25%, which can realize the adaptive extraction of the fault features of the aircraft landing gear hydraulic system, and the features after multi-dimensional fusion have better discrimination, compared with traditional feature extraction methods more effective and more accurate.
The inverse method of characteristics (iMoC) is often employed to simulate the axisymmetric flow behind a predefined shock wave shape. Firstly, two marching schemes of iMoC were compared in this work: the intersection of the left- and right-running Mach lines and the intersection of the left-running Mach line and streamline. The intersection of the left-running Mach line and streamline is discovered to be simpler and more stable. Following that, the applicability of iMoC was analyzed by dividing the shock wave shapes into concave and convex curves, respectively. Based on the oblique shock relations, it is proved mathematically that iMoC is capable of computing the flow behind a concave shock wave. However, for the convex shock wave, if the shock wave angle declines too much along the axial coordinate, the clusters of left-running Mach lines may intersect, leading to iMoC failure. Furthermore, a method for enabling the implementation of iMoC is proposed by generating an expansion fan to replace the infeasible segment of the convex shock wave. Finally, computational fluid dynamics techniques were applied to validate the proposed methods and analysis. This study provides theoretical support to promote the application of iMoC in waverider and inward inlet design.
Aiming at the key scientific problems of strengthening cooling and heat dissipation of turbine blades, a new type of corrugated channel cooling structure was proposed in this work. Refined numerical simulation was carried out in six different corrugated channels, and the physical mechanism of flow and heat transfer was analyzed in detail. The results show that the corrugated channel has smaller flow resistance than the ribbed channel. Among the corrugated channels, the heat transfer effects of sinusoidal, triangular, and sawtooth2 corrugated channels are stronger than rib corrugated channel at the high Reynolds number, and these effects are distributed evenly. Within the six kinds of corrugated channels, the sinusoidal corrugated channel has the best comprehensive heat transfer effect. The blades with sinusoidal corrugated channels have better cooling effect due to the significant increase in heat transfer coefficient on the inner channel surface, and the channel contraction area plays a decisive role in the enhancement of the cooling effect. The heat transfer capacity of the sinusoidal corrugated channel is closely related to the shape of the corrugation, and the heat transfer effect reaches for the best near the waveform H/L = 0.115. The physical mechanism of heat transfer enhancement of the corrugated channel is revealed in this paper, which can provide a reference for the design of the cooling structure of aeroengine blades in the future.
The double-wall cooling as one of the most important cooling techniques of the hot-end components of aeroengine, the heat transfer performance of double-wall cooling structure can be significantly enhanced by the combination of jet impingement cooling and pin fins. Therefore, the double-wall cooling structure with jet impingement holes and pin fins has very important engineering application value. For the double-wall cooling structures with jet impingement holes and pin fins, the experimental studies have been carried out based on various of geometrical and flow parameters. Furthermore, the corresponding numerical simulations have also been carried out according to the corresponding experimental study conditions. In order to approximate to the real double-wall enhanced heat transfer structure of the hot-end components of the aero-engine, in present experimental and numerical studies, the double-wall enhanced heat transfer structure with the double-side outflow channel and the advantage of structural symmetry is adopted. Based on the numerical simulation results, the internal flow field structure, the flow resistance characteristics and the interior enhanced heat transfer performance of the double-wall cooling structure with jet impingement holes and pin fins are analyzed in detail. In addition, the main reasons why the pin fins can enhance the heat transfer performance of the double-wall cooling structure that with jet impingement holes and pin fins are further revealed. Actually, the enhanced heat transfer performance of the double-wall cooling structure is the results of the interactions between the internal jet impingement cooling, the internal crossflow and the turbulent flow of pin fins. Moreover, the influences of the jet impingement Reynolds number (Re), the non-dimensional row spacing (X/D) and the nondimensional hole pitch (Y/D) of jet impingement holes, and the dimensionless diameter (Dp/D) of the pin fins on the enhanced heat transfer performance, the internal flow characteristics, the distributions of local Nusselt number Nu and the flow resistance characteristics of the double-wall cooling structure with jet impingement holes and pin fins are further revealed. Importantly, for all numerical simulation conditions, the distribution differences of the enhanced heat transfer performances on the wall surface of the pin fins and on the wall surface of the jet impingement target plate with the variations of the geometrical parameters and the flow parameters are analyzed in detail. The numerical results indicate that the area-averaged Nusselt number N- u on the wall surface of the pin fins is higher than that on the internal wall surface of the jet impingement target plate. With the increase of the pin fins diameter Dp/D, the N- u on the whole internal wall surface of the double-wall cooling structure presents a tendency of firstly decrease and then gradually increase. In addition, the N- u on the entire interior wall surface of the double-wall cooling structure decreases with the increase of row spacing X/D, but it is not sensitive to the variation of hole pitch Y/D. Moreover, for the discharge coefficient of the double-wall cooling structure, it gradually increases with the increase of the jet impingement Reynolds number Re, the row spacing X/D and hole pitch Y/D of jet impingement holes, but it decreases with the increase of the diameter D-p/D of pin fins. Finally, for 3 <= X/D <= 5, 3 <= Y/D <= 5, 1 <= D-p/D <= 2, and 9000 <= Re <= 30,000, a correlation of the areaaveraged Nusselt number N- u is given, which is in good agreement with the experimental results.
为解决航空并联混合动力系统直接使用传统涡扇发动机时存在的发动机效率下降与低压压气机喘振问题,提出了一种通过质量流量预测确定涵道比的并联混合动力专用涡扇发动机快速设计方法,使用PROOSIS搭建了并联混合动力涡扇发动机模型,对发动机设计结果进行了性能评估与能量利用分析.研究表明,在与基准发动机相同的涡轮前总温限制下,设计结果能够满足推力需求.与在并联混合动力系统中使用基准发动机相比,使用设计的专用发动机时的油耗、能耗、低压压气机防喘振性能更优.混合度越高,使用专用发动机产生的性能提升越大.由于能量利用历程不同,发动机外涵道电能利用率远高于内涵道电能利用率和燃油利用率,这是并联混合动力涡扇发动机节能的根本原因.
为提高用于叶型优化设计ResNet深度学习模型的泛化性,对一种适用于亚声速和跨声速的压气机叶型进行了参数化设计.叶型基于Matlab进行几何模型构建,设计变量为叶型最大厚度、最大厚度位置、栅距和叶型尾缘与轴向的夹角,几何模型通过Pointwise软件进行批量网格划分,网格量级为30万,通过OpenFOAM流体仿真软件进行批量计算.最终通过4个设计变量参数化建模后进行仿真得到了叶型流场仿真数据集,该数据集包含22331个叶型仿真算例,可为ResNet深度学习模型提供训练集和测试题,有助于提高模型的泛化性.
针对某创新构型的涡轮发动机,在国内首次开展大直径小环腔燃烧室的设计及试验研究.提出基于扩压器逆向进气条件下的环涡流场匹配切向燃油喷射的燃烧室设计方案,开展不同状态下燃烧室冷态及热态性能试验,得到燃烧室流阻特性、地面点火特性、贫油熄火特性以及燃烧效率、出口温度场等特性,结果表明:①该设计方案可以满足燃烧室的设计要求;②与常规燃烧室相比,该方案的喷嘴间距比设计达到1.65;③该方案燃烧室的点火性能优异,最低贫油点火油气比达到0.016,点火联焰时间可在4 s以内;④该方案燃烧室的总压损失、燃烧效率、出口温度分布系数(OTDF)、出口径向温度分布系数(RTDF)等综合性能优异,其中慢车状态的燃烧效率能够达到98.6％,设计点的OTDF达到0.16.
建立了一种考虑出口温度分布系数(OTDF)和氮氧化物(NOx)排放特性的航空发动机部件级模型,为发动机燃烧室出口温度分布和排放控制研究提供了仿真平台.以某变循环发动机为研究对象,依据其设计点参数设计燃烧室三维模型,基于CFD数值仿真方法,得到该燃烧室三维模型在海平面不同工作状态下的OTDF特性、氮氧化物排放特性.基于此,建立了适用于全包线、全状态下,可计算燃烧室特性参数的变循环发动机部件级模型.与传统部件级模型相比,该模型能准确地计算发动机在不同工作状态、不同包线点下的燃烧室出口温度分布、氮氧化物排放.仿真结果表明:燃烧室出口温度分布系数与发动机工作状态呈负相关关系,发动机转速越大,OTDF越小,燃烧室出口温度分布品质越好;燃烧室出口氮氧化物排放量与发动机工作状态呈正相关关系,发动机转速越大,燃烧室出口氮氧化物排放量越多,符合发动机燃烧基本规律.
针对传统图像分割方法提取结冰显微图像中的气泡漏检率高和无法分离粘连气泡的问题,提出深度神经网络和传统分割算法相结合的方法.基于Attention U-Net网络,采用双分支融合预测策略对结冰显微图像中的气泡进行提取.针对部分气泡粘连问题,引入直方图均衡化和局部极小值,采用基于距离变换的分水岭算法,对结冰显微图像中黏连气泡进行二次分割.实验结果表明:通过双分支融合预测的AttentionU-Nct网络,对不同结冰显微图像中的气泡提取更精确,特别是对于较小气泡的检出率更高.测试图像的像素精度、平均像素精度、平均交并比和频权交并比分别达到0.976 7、0.891 6、0.818 8和0.957 5.基于距离变换的分水岭算法在粘连气泡分割中也展现了良好的性能,为后续统计气泡个数、面积等特征提供可量化的数据支撑.
针对转子光学压敏测量技术的需求,提出了基于CCD(charge-coupled device)相机的转子短曝光图像增强采集方法,并组建了相应的测量系统,采用两种方法(短曝光图像增强采集与瞬态图像采集)的9个曝光时间对转速为2700 r/min小型风扇转子叶片进行了图像采集.实验结果表明:该转子短曝光增强采集方法和测量系统可有效提升转子图像质量;转子短曝光图像增强采集得到的旋转状态图像模糊长度小于4像素时,实现图像"无运动模糊"采集;采集次数满足静止图像曝光时间与旋转图像曝光时间之比时,旋转状态图像亮度与静止状态图像相当.
为满足对宽速域变后掠飞行器的迫切需求,设计了一种适用于分布式驱动的局部旋转变后掠机翼的过约束冗余驱动机构.以传动性能为指标对单元驱动机构进行尺度初步设计,后结合SQP(sequence quadratic program)算法以机构变形全过程的能量转化率为优化目标对机构尺度进行了优化,优化后的驱动机构在恒定作用力下的输出功提高了 44.3％,能量转化率提高了 37.5％,驱动距离缩短了 9.7％.为解决多个驱动支链驱动力如何分配的问题,将分析超静定结构内力的力法与传统机构受力分析方法结合提出一种准静态驱动力的求解方法,对一定负载及构件材料条件下的四翼梁模型进行了驱动力计算并基于ADAMS(automatic dynamic analysis of mechanical system)在相同负载及构件材料条件下做了动力学仿真实验验证驱动力分配模型的精确度,误差分析显示该模型对于准静态驱动过程驱动力计算误差小于5.5％.最后综合仿真结果及驱动机构的质量对驱动链数目进行优化,确定了最佳驱动链数目为3个.
结合拍振理论仿真分析、实测振动信号分析和工程实际拍振总结,讨论了航空发动机存在的3种多源拍振模式;建立了拍振故障识别流程和排除方法.工程试验验证表明:对于航空发动机常见的3种多源耦合拍振模式,当激振频率相差大于3％时,即可消除多源拍振及其引起的振动值波动问题.所建立的流程和方法在某型发动机振动波动问题排故中进行了应用验证,准确识别出低压转子2倍频-高压转子基频耦合拍振.通过控制规律微调低压转子转速,使低压转子2倍频和高压转子基频相差3％,消除了拍振引起的振动波动,验证了研究结论的正确性.
翼吊螺旋桨发动机的短舱对飞机的升力失速特性有明显的影响,采用数值模拟方法研究了短舱扰流片对失速的影响效果.数值仿真和风洞试验表明,在零推力状态,某大型四发螺旋桨飞机在超过失速迎角以后,内外发动机之间的机翼首先分离,并快速推进到机翼前缘,失速以后升力损失达到最大升力系数的30％左右.为了限制机翼的分离速度,在外发短舱的内侧,安装了一个扰流片.仿真结果表明,在最佳设计位置,明显改善了失速特性,与无扰流片状态相比失速以后升力损失减小50％左右,失速迎角没有明显变化.不同的扰流片安装位置对失速的改善效果差异明显,从最佳位置向周向以及螺旋桨方向移动会造成扰流片失效,向机翼方向移动对改善机翼分离速度有效,但是会明显降低失速迎角,向机翼方向移动后,受周向位置的影响减小.针对最佳设计位置开展了着陆构型零推力状态的风洞试验验证,加装扰流片以后,机翼失速后的升力损失由0.92降低至0.42左右,升力损失减少54％,与数值仿真结论基本一致.
针对文献中采用离心喷嘴的氢氧预燃室热试车出现液膜声学相关频率的现象,采用两相流温度瞬态传热仿真计算了 30s热试车中喷嘴内液氧膜平均温度的变化过程.为节约计算资源和时间,采用Flu-ent 多相流模型计算获得稳态等温速度场,采用二维轴对称传热模型调用两相流速度场和相分布结果进行流固耦合瞬态温度仿真.传热计算结果显示:喷嘴出口燃气温度对液氧膜温度有较为显著的影响.喷嘴出口气体温度升高50 K会导致液氧平均温度升高约2K,喷嘴出口气体温度为105 K时,仿真计算得到的液氧平均温度数值与文献中经验修正值较为吻合,验证了其温度经验修正方法具有一定的合理性.模型计算结果能够解释固体热容对液氧加温作用导致喷嘴声学频率降低的现象.
以某型号二冲程航空活塞发动机为研究对象,通过建立一维、三维发动机模型、喷油器模型和空燃比控制模型,辨析进气量和喷油量的主要影响参数,基于递归神经网络进气量预测和喷油模型,在变工况下对二冲程发动机空燃比控制进行研究.在节气门开度变化的瞬态工况下,将空燃比控制的超调量控制在4.6％以内,能在变工况停止后较短的时间0.3 s内将缸内混合气恢复至当量比.在不同海拔工况的仿真研究下,随着海拔的增高,空燃比控制模型的超调量和回调时间适当减小并逐渐稳定.
针对双转子系统低压激励和高压激励振动模态的正交性受转速比影响,提出了以定转速比转子模态替代实际工作转速线下的转子模态进行双转子系统N1+N2和N1+N2+4平面模态动平衡的方法.推导了N1+N2和N1+N2+4平面模态动平衡法的平衡条件,给出了模态不平衡质量和正交校正质量组的计算式,并以某型航空发动机双转子系统为例,给出了N1+N2模态动平衡方法平衡各阶模态的过程.最后在双转子实验系统上进行了动平衡验证实验,实验研究发现在定转速比的条件下动平衡的减振效果最大可达72.4％;当高、低压转子转速控制率为实际工作转速线时,在工作转速范围内,各阶临界转速处的振动幅值总量都降低到120 μm以下,满足设计要求.实验结果表明提出的双转子航空发动机模态动平衡方法是可行的.
基于金属橡胶减振器实际服役场景,设计了金属橡胶减振器横向和轴向准静态加载夹具,获取了 5种型号共175个金属减振器横向和轴向准静态力学性能数据.借助数据分析软件对试验数据进行高阶多项式拟合,成功确定了 5种金属橡胶减振器的横向和轴向载荷-位移曲线平均值和性能边界.结果表明:减振器横向和轴向载荷-位移曲线分散性并不服从常见的高斯概率分布规律,且难以同时兼顾,轴向加载时MR2型减振器分散性最小,MR3型减振器分散性最大,而横向加载时MR3分散性最小,MR2分散性最大.相关性能边界可为金属橡胶减振器的选型和安全性、可靠性评定提供重要支撑.基于μCT三维断层扫描分析设备,亦建立了金属减振器中金属丝组件的三维编织形貌,并成功确定了不同高度区域金属丝体积.
为提高燃机过渡态总体性能仿真精度,采用模块化思想建立和完善了二次空气系统精细化模型,提出了燃机总体性能与二次空气系统(SAS)耦合的过渡态仿真方法.以双轴燃机为研究对象建立了仿真算例,模拟并分析了突增、突卸和突甩负荷时二次空气系统各支路引气和汇流的动态变化对总体性能参数的影响.结果表明:通过耦合仿真可以评估燃机运行工况的变化对引气比的作用效果,在突增、突卸负荷时,二次空气系统引气比变化量为0.18％,而对于突甩负荷这种极端工况,引气比的变化量增大至0.55％;燃机主流道与二次空气系统之间的动态交互作用不会显著影响过渡态总体性能参数,但对二次空气系统各支路非均衡响应过程有较大影响,这是现代燃机精细化仿真中不可忽略的因素.
